Integrated vehicle fluids

ABSTRACT

A system and methods are disclosed for an upper stage space launch vehicle that uses gases from the propellant tanks to power an internal combustion engine that produces mechanical power for driving other components including a generator for generation of electrical current for operating compressors and fluid pumps and for charging batteries. These components and others comprise a thermodynamic system from which system enthalpy may be leveraged by extracting and moving heat to increase the efficient use of propellant and the longevity and performance of the launch vehicle.

CROSS-REFERENCE TO RELATED APPLICATION

This application is a continuation of U.S. application Ser. No.16/111,623, entitled “Integrated Vehicle Fluids”, filed on Aug. 24,2018, now U.S. Pat. No. 11,261,828, which is a continuation of U.S.application Ser. No. 15/795,675, entitled “Integrated Vehicle Fluids”,filed on Oct. 27, 2017, now U.S. Pat. No. 10,718,294, the entiredisclosures of which are incorporated herein by reference in theirentireties.

FIELD OF THE INVENTION

The invention generally relates to rocket propulsion systems for spacelaunch vehicles placed and maintained in planetary orbits. Morespecifically, the invention relates to systems and methods forpropelling and powering an upper stage of a space launch vehicle byusing ullage gases generated by the propellants to generate power by aninternal combustion engine. The power generated by the internalcombustion engine replaces power and reaction control systemstraditionally powered by separate hydrazine systems and batteries, andeliminates the need for separate systems used for pressurizing the mainpropellant tanks, such as helium-based systems.

BACKGROUND OF THE INVENTION

There are increasing demands to lower the cost of space transport togeostationary and other orbits as well as perform missions which aresimply not possible with existing launchers such as manned explorationof the moon and Mars. Implicit is the demand that reliability beenhanced and certainly not degraded. Unspoken but also implicit is thatcommercially viable launchers must fill this broad range of demandssince single-purpose launchers such as the Saturn rocket or SpaceShuttle are cost prohibitive, even for governments with substantialspace budgets. The commercial launchers presently being used for mostmissions are the result of decades of evolution and have become highlyrefined and proven. Each individual system on the launchers has beenraised to a high level of performance which is very difficult to improveupon, even with large investments in engineering development. Since onlyincremental improvements can be expected by addressing individualsystems, there is a need to view space vehicle systems in a more broadsense to determine if substantial improvements can be accomplished.

One example of a high performance, highly evolved upper stage is theCentaur®. The Centaur® upper stage is capable of delivering payloads tonearly any orbit from Low earth Orbit to interplanetary trajectories.The Centaur® is capable of delivering a high energy to the payload byburning liquid hydrogen (LH₂) and liquid oxygen (LO₂) in a very highefficiency, low weight engine such as the RL10. The total mass of theCentaur when empty is less than 2.5 mt, yet it can contain in excess 20mt of propellant. Propellants are stored in lightweight stainless steeltanks whose structural rigidity is provided primarily by the pressure ofthe propellants within the tank. In order to keep the tanks fromcollapsing prior to the loading of propellant, the tanks are pressurizedwith gas. In the Centaur®, a common double bulkhead is used to separatethe LO₂ and LH₂ tanks. The two stainless steel containers are separatedby a very thin layer of insulator which is contained within a hermeticcavity. Therefore, the appearance is of a single tank, but it is dividedinto the separate LO₂ and LH₂ tanks with an intermediate vacuum cavity.The extreme cold of the liquid hydrogen on one side of the bulk headcreates a vacuum within the intermediate cavity. The insulator preventsthe two metal bulkheads from contacting thereby maintaining a lowthermal conductivity, thus preventing heat transfer from thecomparatively warm LO₂ to the super cold LH₂. The exterior of the tanksare also insulated to suppress heat flows from the external environmentto the propellants.

While on the ground and filled with propellants, the tank pressures arecontrolled by valving which maintains the tanks within a specificpressure band. The propellants within the tanks boil due to externalheating and the vapor formed is passed through these regulating valveswhich hold the tank internal pressure within band regardless of theheating and attendant boil off vapor mass flow. By controlling the tankpressure at which the propellants boil, their saturation conditions areestablished. For the sake of maximizing the density of the propellantsand hence reducing the size of the vehicle tanks, the pressures andtemperatures are kept as low as possible within the tanks. These ventvalves are thus precision cryogenic regulators that are complex, costlyand heavy.

While on the ground, loads imparted to the vehicle are quite low, andthe stiffening effects of the low internal pressures controlled by ventvalves are sufficient to maintain structural integrity of the vehicle.However, during the ascent phase of flight and also prior to operationof the upper stage engines, the pressures within the vehicle tanks mustbe raised. In the case of ascent, the vehicle must be further stiffenedso that it can survive the very high bending and compressive loadsgenerated by aerodynamic, thrust and inertial effects. Pressures areraised prior to engine start to permit the proper operation of theengine pumps. These high capacity pumps must receive propellants whosepressure is substantially above their saturation pressure. Thissaturation pressure was effectively set prior to liftoff by the valvingcontrolling tank pressures. Without system pressure maintained abovesaturation pressure, the propellants would boil within the pumps andthey would cease to function properly. This margin is commonly referredto as Net Positive Suction Pressure (NPSP) and is commonly on the orderof 3-10 psi.

In most modern upper stage vehicles, these pressurization demands aremet by introducing gaseous helium into the ullage spaces of thepropellant tanks. This helium is stored in separate vessels, typicallyat high pressure, and is delivered via valves to the propellant tanks atneed. Helium is used since it has a low density, is chemically inert,and does not condense to a liquid at the cryogenic temperatures seen inthe LO2 and LH2 tanks. Hence it can be used to pressurize both the LO2and LH2 tanks with a mass penalty. Once the upper stage engines areoperating, it is possible to perform the pressurization task by bleedingsmall amounts of warm H2 and O2 gases from the engine. This reduces theamount of helium required for the mission. The amount of helium requiredis thus dictated by the size of the propellant tanks, their pressure andthe number of burns which are expected to be performed. The mass of thehardware required to contain this helium is very significant and manyapproaches have been taken to suppress system complexity and weight.However even the most advanced existing systems have strict limitationson their capabilities. These systems all have a limited amount of GHeand hence the number of engine burns, tank size, and other factors areall limited. Even a small leak of helium from the storage systems canresult in a catastrophic loss of pressurant and hence mission failure.

During flight the upper stage propellant tanks will continue to absorbenergy from the environment, albeit at a lower rate than what waspresent prior to launch. During engine burns, elevated tank pressuresare maintained with GHe, gaseous O2 or H2 to establish and maintainsufficient NPSP and hence will end up at the end of a burn at a tankpressure above the saturation condition of the propellants. As heat isapplied to the liquid propellants, they will gradually increase intemperature until their saturation pressure matches the partial pressureof H2 or O2 in the ullage gas. At this point, the propellants begin toboil. Tank pressures rise as the boil-off continues. If no action istaken prior to the next start of the engines, the system must bepressurized above this new higher pressure. The incremental increases intank pressures therefore directly drive the peak operating pressures ofthe tanks, and hence their mass. Therefore, tank designs may have toaccount for much higher pressures, such as a 60 psia capability, whichresults in a substantial mass penalty.

To mitigate this pressure building effect on missions lasting more thana few minutes, it is common to vent the pressure in the ullage space toa level close to the original saturation pressure. Especially on the LH2tank, during a mission to geostationary orbit, this venting may beperformed multiple times. The energy absorbed from the environment isstored in the enthalpy of the ullage gases which therefore must besubsequently dumped overboard.

A significant limiting factor for all missions in space is accountingfor the ullage losses associated with the continual boil off ofcryogens. It is this propellant loss that has prevented to date the useof cryogenic propulsion systems for missions to the moon or indeed anymission with a duration that is much longer than one day. One of themost effective approaches for reducing losses is to apply a very lowthrust to settle the propellants within the tanks to fixed locations,generally towards the aft end of their respective tanks. Less than athousandth of one G is required to achieve this effect. Settling thrustsegregates the liquid and gaseous phases of each propellant. Cold liquidpropellant is thus physically separated from much warmer gas by thesettling thrust. The quiescent gaseous ullage, in a microgravityenvironment without significant convection, then behaves as an excellentinsulator and blocks heat from entering the liquid propellant surfaces.Heat is conducted down the side walls from the warm ullage side of thetank to the cold liquid side but this is inhibited by the relativelylong conductive distances, reduced thermal conductivity due to thecryogenic operating temperatures and low wall thicknesses. Naturally thethinner these walls are the better. Thus it can be seen that a tank witha low gage and hence low allowable operating pressure is also thermallysuperior. All of these effects conspire to slow boil off when settlingis imposed.

Settling thrust is typically provided by one or more small rocketsfueled by hydrazine. On the Saturn S-IVB stage, the ullage gases wereburned in a small thruster to maintain vehicle settling and some of theheat of the burning H2 and O2 was used to warm cold helium up for use inthe pressurization system. Other vehicles such as the Delta CryogenicSecond Stage simply vent the boil-off gas aft to produce a small amountof settling thrust during long duration missions. Most often though,these waste ullage gases are simply dumped. Depending on vehicle designand mission duration, these wasted propellants can weigh into thehundreds of pounds. Naturally, the amount of time that a hydrazinesystem can support settling is strictly limited by the amount ofpropellant that it contains. Despite the best conservation efforts,hydrazine-based settling can at best be sustained for a handful ofhours. Once settling is lost the surface tension effects within thepropellants will gradually cause the interior of the propellant tanks tobe fully wetted, temperature segregation will be lost and boil off ratescan triple.

The vehicle must also provide a means for changing its attitude, andthis function is also typically done with a hydrazine fueled thrustersystem. On the Centaur® vehicle, the attitude control thrusters andsettling thrusters share a common supply system. While the settlingfunction consumes the vast majority of hydrazine capacity, the attitudecontrol task cannot be ignored. Settling thrusters can be commanded off,but the vehicle attitude must be stabilized for various reasons toinclude (i) maintaining radio links to the ground, (ii) providing anoptimal attitude relative to the sun so that components such as avionicsdo not get too hot or cold, and (ii) suppressing heating of the mainpropellant tanks. Even if settling is eliminated, the attitude controlfunction alone can consume hundreds of pounds of propellant over thecourse of a multiday mission. This propellant requirement isinsupportable by most commercial upper stage launch vehicles.

Regarding the use of hydrazine as a propellant, while its application tospace vehicles is widespread, there are a number of problems associatedwith its use. Hydrazine is a highly toxic, highly corrosive fluid thatis compatible with only a handful of materials. Handling hydrazinerequires hazardous procedure precautions, often requiring the use ofpositive-internal pressure inflatable SCAPE (Self Contained AtmosphericProtective Ensemble) suits to protect technicians loading a vehicle.Hydrazine can only be used in a narrow band of temperatures near roomtemperature. Hence, elaborate thermal control measures including heatersare mandatory, thus burdening the electrical storage system andexacerbating propellant heating. Hydrazine is also quite costly.Hydrazine is also a very inefficient fuel, delivering only a miserlyspecific impulse of 235 seconds. The advantages of Hydrazine are that asa fuel, it is simple and reliable to use assuming the appropriateenvironmental conditions can be maintained during its storage anddelivery to a reaction chamber. Hydrazine is catalytically decomposed ina simple reaction chamber and does not require an ignition system oreven an oxidizer. Nevertheless, the continued use of hydrazine setsharsh boundaries on improving overall vehicle operations and costs.

The electrical systems on the upper stage currently use largeelectrochemical batteries to provide power. This battery technology hasevolved over decades to favor batteries of increasing power density andattendant sophistication. The desire to provide redundancy has doubledthe storage demand. Even with the best modern technology, thesebatteries are extremely heavy, costly and can only supply enough powerfor less than a day's operation of a vehicle such as a Centaur®. Withouta means to recharge these batteries, they set a strict limit on missionduration. Unfortunately the two common sources of power for recharge aresolar panels and fuel cells, and these systems are both very costly toincorporate on a vehicle. Use of solar panels requires vehicleorientation control relative to the sun, and are physically bulky withcomplex deployment mechanisms. Most spacecraft that use solar panels areeffectively in zero-G conditions, and hence large deployed solar panelsare never exposed to high loads. A vehicle like Centaur® will generateacceleration forces in excess of 2 G's, and hence the mounting systemfor even a small solar array would be very heavy.

Fuel cells, while being more compact than batteries and are efficientand seemingly simple, are quite costly and complex to operate andsupport due to their intolerance of inert gases within the reactantstreams and due to the necessity to dispose of the water they produce.To date, only manned vehicles such as the space shuttle can justifytheir cost and complexity.

While cost reduction, increased simplicity and reliability are primarygoals in an improved vehicle, there is also an increasing need to expandmission capabilities beyond merely moving heavier payloads. Currentmissions are performed over a maximum flight duration of less than aday. However if the vehicle could efficiently fly for longer, it wouldbe extremely valuable. Missions such as those to the moon require coastdurations measured in days. The increasing amount of space junk in orbitwill soon require the deliberate disposal of not only obsoletesatellites but also the stages which placed them in orbit. This disposalactivity at present would impose large performance penalties which woulddrastically increase the cost to orbit. However by performing disposalmaneuvers at optimal times, this function can be accomplished with aminimum of cost. Missions such as space junk removal require a vehiclebe capable of flying for days to weeks. In summary if one wishes toimprove vehicle system performance and cost, yet expand the missionduration and improve reliability, a broader view of the vehicle must betaken to include a simultaneous analysis of vehicle thermodynamics,power, propellant and pressurant storage limitations, vehicle structuraland thermal interactions, and the demands of widely varying missions.While it may be possible to redesign vehicle systems on a micro level,that is, to redesign selected systems based on specific missionrequirements, this design approach inevitably compromises the majorityof missions and can also create a proliferation of system designs thatare all slightly different and likely incompatible. This micro leveldesign solution is the origin of the present state of most space vehiclecapabilities.

The use of waste ullage gas was recognized in the 1960's as a potentialsource of fuel for an auxiliary engine on the Saturn S-IVB. NASArecognized that these ullage gases could be captured and reused withinan internal combustion engine that could be used to provide power forthe upper stage vehicle. Although this recycling or reuse of the ullagegases was recognized development stalled with the proof of concept of aH2/O2 burning internal combustion engine. The concept was never flown.

There are a number of examples of improvements made to rocket propulsionsystems in order to increase main engine propulsion efficiency, or tosimplify the components of a launch vehicle, with one intent beingincreasing the available payload of the vehicle.

One example of such a reference is the U.S. Pat. No. 5,282,357 for ahigh-performance dual-mode integral propulsion system. This referencediscloses a propulsion system in which pure hydrazine is used as thefuel for both a bi-propellant rocket engine for high thrust performanceand in multiple mono-propellant thrusters for station keeping andattitude control functions. The use of the common fuel for both modes ofoperation significantly reduces propellant weight and inert propulsionsystem weight. For station keeping, the mono propellant thrusters can beaugmented in performance by employing either electrothermal oradditional direct chemical energy, arc jet operation, or force fuelacceleration to provide increased specific impulse values.

The U.S. Pat. No. 6,135,393 provides a spacecraft attitude and velocitycontrol thruster system that incorporates mono-propellant RCS thrustersfor attitude control and bi-propellant scat thrusters for velocitycontrol. Both sets of thrusters are designed to use the same liquidfuel, supplied by a pressurized non-pressure regulated tank, and operatein a blow down mode. In the propulsion system, such station keeping andattitude control thrusters may function in conjunction with a largethrust apogee kick engine that uses the same propellant fuel. Hydrazineand bi-nitrogen tetroxide are preferred as the fuel and oxidizer,respectfully.

Despite improvements in general rocket technology, to include increasingthe efficiencies of rocket engines and components, there is still a needto provide even greater efficiencies, and to simplify space launchvehicle systems while carrying larger payloads over longer durations.

SUMMARY OF THE INVENTION

In accordance with aspects of the present disclosure, a system, methodsand sub-systems or sub-combinations are provided to supply all requiredvehicle functions including attitude control, propellant settling,propellant tank pressurization control, hardware actuation and purgingand power generation without the need for any fluids other than theprimary propellants present in the main vehicle tanks. For simplicity,LO2 and LH2 will be described herein as the primary propellants.However, it should be appreciated that other liquid propellants may besubstituted. For example, according to aspects of the presentdisclosure, the oxygen and/or hydrogen may be replaced with methane,kerosene, nitrogen tetroxide, hydrazine, hydrogen peroxide, nitrousoxide, etc. The system thus has a capability to perform these functionswithout any limits other than the mass of the primary vehiclepropellants. The overall system can be referred to as an integratedvehicle fluid (IVF) module that provides these functions. The waste orullage gases from the hydrogen and oxygen tanks that are typicallyvented overboard, are used as the fuel and oxidizer to run an internalcombustion engine to provide power for other vehicle functions. Thepower output shaft from at least one engine can be used to drive anumber of secondary devices to include one or more generators thatgenerate electrical energy for storage in one or more small batteries.Power from the shaft is also used in other secondary devices such as oneor more pumps to modulate and regulate fluid pressures in the vehicle,to include most importantly, pressures in the LO2 and LH2 tanks.Alternatively or additionally, power from the batteries may be used todrive the various components.

In a first exemplary embodiment, the IVF module has one or more pumpsthat move a substantially liquid propellant using mechanical work andthat generate a large pressure differential to operate variouscomponents, including accumulators. The one or more pumps circulate thepropellant through heat exchangers that transfer heat directly fromthrusters to the propellant. Similarly, propellant such as hydrogen gasis circulated around the internal combustion engine to transfer heatdirectly from the internal combustion engine to the hydrogen gas, whichis then fed into the internal combustion engine.

In a second exemplary embodiment, the IVF module has one or more pumpsthat move a substantially gaseous propellant using mechanical work andthat generate a small pressure differential. A separate coolant loopscavenges heat from various components of the IVF module and transfersheat to the propellants through heat exchangers. Aspects of these twoembodiments are described in more detail below. It will be appreciatedthat these embodiments are exemplary in nature and that one aspect fromone embodiment could apply to the other embodiment, or any otherembodiment. For instance, the type of engine used in one embodimentcould be used in the other embodiment.

Regarding the first exemplary embodiment, one or more internalcombustion engines (ICE) can be one of many selected types of engines toinclude a piston engine or a Wankel engine. This engine burns thegaseous hydrogen (GH2) and gaseous oxygen (GO2) from the upper stagepropellant tanks. The GH2 is introduced into the engine through a flowcontrol valve that throttles the mass flow of the GH2. Before the GH2enters the engine for combustion purposes, the hydrogen is used to coolthe exterior of the engine, maintain pressure in the crank case, andcool the internal chamber of the engine. The GO2 is injected either intothe hydrogen before or at the intake port or directly into the enginecombustion chamber at an elevated pressure, similar to fuel injection ina diesel engine. In circumstances when the space vehicle requiresadditional power output from the engine, it is also contemplated thatLH2 can be mixed with the GH2 bled from the hydrogen tank to temporarilyfeed the engine. This additional cold fluid enables greater coolingcapacity for the engine and increases the H2 density within thecombustion chamber thus allowing more O2 to be introduced and hence moreenergy released. This mixing can be modulated by an intake controlvalve. Under most circumstances, however, the ullage hydrogen is allthat is required to power the engine in order to provide sustained powerfor all upper stage systems.

In another important aspect of embodiments of the present disclosure,the exhaust from an internal combustion engine, composed primarily ofhigh temperature hydrogen rich gas, is used to power one or more of thesettling thrusters with a very high efficiency of thrust generation, asdiscussed further below. Therefore, the hydrogen ullage gas undergoes asingle recycling use through the engine. Either a piston engine, Wankelengine, or turbine engine is contemplated for use in the presentinvention. As compared to a piston engine, a Wankel engine has no intakeor exhaust valves, has fewer moving and lubricated parts, a very simplecrank system, and is a dynamically balanced engine due to itssymmetrical disposition with respect to the movement of the rotor.Because of this simplicity the engine is very light. The three chamberconfiguration of the rotary engine has distinct hot and cold areas thatcan enable very simple gas cooling, yet may achieve higher exhaust gastemperatures that can be used as the exhaust gas for the settlingthrusters.

Regardless of the type of engine used, because of the limitations oncombustion physics and materials, the engine preferably operates at avery low mixture ratio between 0.6 and 2. This range precisely matchesthe boil off characteristics of the vehicle which often generates moreH2 than O2. Hence the engine more effectively uses the waste gases fromthe vehicle. According to some embodiments of the present disclosure, toaccomplish all known vehicle functions requires a total shaft power ofless than 10 kW and more commonly less than 2 kW. This allows the engineto be only moderately efficient and with a very small displacement onthe order of 200 cc. High exhaust pressures can be tolerated by theengine, (such as in the range of 10-20 psia) which enables at least a5-10 psia thruster chamber pressure in the settling thrusters by simplydirect venting from the engine exhaust into the receiving chamber to thethrusters, without any further pressurization requirements. Settlingthrust thereby generated is in the precise band to provide continuouslow G forces. With respect to generation of electrical power, a verysimple yet effective electric power generation is achieved by anelectric starter/generator that is driven by the output shaft of anengine. The starter/generator generates electrical current for storagein one or more rechargeable batteries. For example, a combination orarray of multiple lithium ion batteries may comprise a single batterypack. The starter/generator, associated electronics and the battery packmay also be cooled with hydrogen flowing through the IVF module. Thebattery pack can then provide electrical power for all the other upperstage systems, as well as power for pressurization pumps to pressurizethe propellant tanks. The battery pack maintains a minimum charge andwill discharge during peak loading conditions. The battery pack iseasily recharged during vehicle coasts in which power loading is reducedsince system duty cycle for lateral thrusters and pressurization is lowduring this period. By inclusion of the battery, this removes manyrestrictions on peak power and total available energy that were aninherent concern for prior space launch vehicles that solely relied uponbattery power for many functions. The battery pack can be less than 5%of the capacity of present batteries with a proportional reduction inmass and volume. The drastic reduction enables the use of less exotic,lower power density battery designs without significant mass penalty butwith large cost benefits.

The starter/generator permits the repeated startup and shutdown of theIC engine as required for ground testing and flight operations. Ifdesired, the engine can be shut down for extended periods and allvehicle power provided by the vehicle battery pack. Settling cancontinue to be supported without the operation of the IC engine andhence low-boil off, settled thermodynamic conditions can be maintainedfor even longer periods. When the batteries have been discharged totheir low limit the IC engine can be restarted and the system loadstransferred effectively to the generator which simultaneously rechargesthe battery pack. In a preferred embodiment the starter and generatorfunctions are performed by a single electromechanical device.

In a preferred embodiment, the IVF system performs the function ofground venting of the main propellant tanks via a ground vent valve.These valves are connected to the airborne system via disconnects whichactuate once the vehicle has achieved liftoff. Because the ground valvesare not restricted in mass, their regulation capability is far moreprecise and stable than a mass and volume restricted airborne valve.Their flow capacity is also far higher and hence higher heating rates orlower tank pressures can be readily accommodated. Since the high heatingrates requiring these valves are not present during flight, the lowercapacity airborne systems are used once the vehicle is placed in thevacuum of space.

One important aspect of embodiments of the invention is that the launchvehicle can operate for long settled periods with thrust supplied onlyfrom the ullage gases. This sustained vehicle settling drasticallyreduces propellant losses in the tanks. Empirical testing has shown thatboil off can be reduced to less than a third of normal losses. Unlikehydrazine based systems, this settling capability is generated by fluidsthat would typically have been merely dumped.

According to aspects of the present disclosure, multiple levels andsources of settling thrust are provided and integrate the airborne maintank vent function into the settling thrusters. Ultra-low forces can begenerated by simply venting cold GH2 gases though the settling thruster.This also accomplishes the venting of the main LH2 tank during flightdescribed in the background. Because this is accomplished with multipleaxial thrusters, the vent rate can be modulated and there are redundantvalves to enable this critical function. Thus a rapid tank vent can beaccomplished without the need for dedicated vent valves as is therequirement with presently known vehicles. These vent events produceaxial thrust which helps further settle the vehicle, and by modulatingthe axial thrusters, the vehicle attitude can be maintained by theguidance system. The need for precisely balanced vent systems asrequired in known vehicles is thus eliminated.

Although a rapid tank vent can be commanded, it may be more effective tosimply burn off the excess ullage. Low thrust for sustained coastsettling, (such as in the range between about 0.5. to 2 lbf) may beprovided directly by the hydrogen rich ICE exhaust gas fed to the axialthrusters. Since the hydrogen fed to the ICE is supplied by the tankullage, this gradually consumes the H2 ullage gas which would have to bevented in any event to reduce tank pressure. According to aspects of thepresent disclosure, combinations of direct H2 tank syphoning and ICEhydrogen burn off can be executed as required by mission needs, externalheating requirements, or power demands. In nearly every case exceptdirect venting of cold gas, the specific impulse of these settlingthrusters are at least 50% higher than for existing hydrazine thrustersthereby providing a powerful performance enhancement.

Greater thrust and gradual LO2 tank vent down is obtained from the axialthrusters by adding GO2 to the ICE exhaust gases at the axial thrustersor in the exhaust lines leading to them thereby increasing the energyrelease. According to embodiments of the present disclosure, by addingGO2 to the ICE exhaust, a demand for GO2 is created from the accumulatorwhich during a coast phase is replenished from the LO2 tank ullage. Thusin addition to direct venting of the LO2 tank through an axial thruster,excess GO2 is effectively burned off via the axial settler.

High axial thrust (4-25 lbf) and simultaneous LH2 and LO2 tank vent downmay be obtained by adding further GH2 from the ullage to the ICE exhauststream as well as adding GO2. This mode is effectively the highest rateof tank venting while generating peak axial thrust. This mode may beused during the highest settling demand periods immediately prior to andafter main engine operation.

The ability to allow tank pressures to be reduced while taking fulladvantage of the vented gas allows efficient re-saturation of the liquidpropellants at lower pressures. Since tank pressures are not allowed torise uncontrolled, design pressures can be reduced on the vehicle maintanks. As described, this design pressure reduction has a profoundpositive effect on vehicle tank mass and its overall thermal efficiency.Both of these factors amplify the performance benefits of the IVFsystem.

According to some aspects of the present disclosure, the lateralthrusters are supplied from small accumulators which are held by acontrol system at a low pressure (approximately 200-500 psia) andnear-ambient temperature. These thrusters burn H2 and O2 at a moderatemixture ratio between 1 and 4. The inlet mixture ratio is bounded by thetemperature and pressure limits imposed on the storage accumulators.Because of this the high performance inherent (a specific impulse inexcess of 350 seconds) in a H2/O2 thruster is obtained without need tobring cryogenic liquids and maintain them in thermodynamic stasis ateach thruster inlet valve. In addition the seals, seats and other softgoods are not exposed to cryogenic conditions and hence are simpler,more reliable and less prone to leakage. Regenerative cooling of thethruster combustion chamber enables the total deletion of hightemperature alloys and complex assembly methods from the thruster.

In some embodiments, the lateral thrusters can operate in two modes. Thefirst is the standard combustion mode whereby H2 and O2 are ignitedwithin the thruster to produce on the order of 10-35 lbf of thrust. Ineach axis there are twin redundant thrusters so two levels of force areimmediately available. The thrusters can also be operated in cold gasmode by only commanding a H2 inlet valve open. This provides acapability to produce very small impulses using a low-temperature andnon-condensable exhaust. Precision vehicle maneuver in close proximityto other vehicles becomes straightforward without the threat of hightemperature plumes either damaging or contaminating sensitive radiationshielding or other elements of the vehicle being docked.

According to some aspects of the present disclosure, the IVF module mayinclude two small accumulators for containing GO2 and GH2. Theseaccumulators are periodically replenished from either the gaseous ullageor from the liquid propellants in the vehicle tanks. Whenever the mainupper stage engines are operating, GH2 and/or LO2 can be bled from themain engine pumps. During coast phases when the engines are notoperating low pressure fluids from the main propellant tanks (eithergaseous or liquid) are pumped up to an accumulator pressure of between200 and 500 psia with small pumps which are driven through clutches orvia electric motors by the internal combustion engine. Fluids exitingthe main engine bleeds or the IVF pumps may be quite cold and mightrequire warming prior to storage in the accumulators. This isaccomplished by warming them in heat exchangers which are part of theexhaust system of the internal combustion engine and the downstreamaxial thrusters. The temperature of these gases is controlled by eithersimple mechanical thermostatic devices or via sensors and activecomputer control. These control devices modulate the amount of heatwhich is added to the cold gases exiting the pumps to achieve a steadytemperature of gas delivered to the accumulators. Gases stored in theaccumulators are thus stabilized within a narrow pressure andtemperature band which is close to room temperature.

During low duty cycle periods the accumulators are replenished via theIVF pumps using ullage gas. This is typical of coast periods of amission when the main engines are not operating and tank pressurizationevents are not occurring. This enables the best use of the waste boiloff gases. Compression of gases though requires a larger expenditure ofenergy by the IC engine due to the larger enthalpy change associatedwith gaseous compression. This limits the total mass flow which can besupported by ullage gas compression. When ullage gases are flowingthrough the pumps, the need for heat addition from the IC engine exhaustand axial thrusters is low since much heating is accomplished simply bythe heat of compression added in the IVF pump.

For high duty cycle periods when rapid and simultaneous pressurizationof both hydrogen and oxygen tanks is required along with high settlingthrust, the pumps consume liquid cryogens which are more efficient tocompress and raise to accumulator pressure. These liquid cryogensrequire more heat addition from the axial thrusters but far larger massflows can be supported by the IC engine power output. The use of liquidcryogens of course debits the vehicle main propellants and decreases theamount of LH2 and LO2 available to the main engines. The amount ofliquid cryogens thus consumed however is compensated by the eliminationof dry mass, hydrazine and helium from the vehicle. Effectively, theseliquid propellants are converted to gaseous pressurants which are thenlater reused to generate power, settling and axial thrust.

The stabilization of the accumulator pressure and temperature simplifiesthe design of downstream devices such as the thrusters andpressurization valves. The thrusters can be operated with a narrow rangeof mixture ratios and thrust output since the inlet conditions arebounded. Similarly the pressurization valves can be sized to addressonly a restricted inlet density band unlike the situation with typicalexisting systems where valves must be capable of throttling inlet gaseswith a pressure band in the thousands of psi and temperature swings inthe hundreds of degrees. The near-ambient storage conditions in theaccumulators also enable the use of elastomeric and other materials inthe construction of downstream valves and components. The combination oflow pressure and ambient temperatures enables leakage of hydrogen andoxygen to be minimized with simple and reliable seats. The necessity forspecialized, low-rate and hence costly cryogenic components is thuseliminated.

According to some aspects of the present disclosure, the oxygen andhydrogen pumps for the IVF module may be small, with displacementstypically on the order of 1-10 cc in some embodiments. The pumps caneither be driven with mechanical clutches off of the internal combustionengine shaft or be driven by motors supplied with electricity from thestarter/generator/batteries. The pumps are commanded by the IVFcontroller to turn on whenever their respective accumulator reaches itslow pressure limit or can be directly commanded whenever significantfluid loads are imminent.

In a fashion similar to the thrusters, according to some aspects of thepresent disclosure, the modules may contain valves whose function is todeliver warm GO2 and GH2 to the respective main propellant tank ullagesfor tank pressurization. These gases are bled from the IVF accumulatorsat need, typically immediately prior to upper stage engine start andduring engine operation but also during booster ascent. The linesleading to the main tanks for pressurization are in one embodimentseparate from the vent lines leading from those same ullage spaces tothe intakes of the pumps and internal combustion engine. This preventsthe ingestion of warm, high pressure gas into these devices which areoptimized to induct the colder, lower pressure gas resident in theullage spaces.

The mass of the ullage gases which remain inside the vehicle at thecompletion of the mission may also dramatically be reduced by the IVFmodule. According to aspects of the present disclosure, the hydrogentank, pressurized with warm H2 from the accumulators, will haveapproximately half the mass as would be encountered in prior systems.This is due to the elimination of GHe (a heavier molecule than H2), theoverall warmer temperature of the ullage, plus the reduction in the peakpressure required. The GO2 ullage is also considerably lighter due tothe increased temperature and decreased pressure.

The IVF system also can supply gaseous H2 and O2 to actuate valves onthe main vehicle and main engines. The propellant flow control systemson the vehicle are often actuated by medium pressure gaseous helium(GHe). Typically this gas enters into a piston in cylinder arrangementand the supplied pressure forces the piston to move which then actuatesa ball, butterfly or poppet valve. The GHe is trapped in a dead-headedcavity adjacent to extremely cold liquid propellants. The gradualcooling of this actuation gas can adversely affect the opening andclosing characteristics of the cryogenic valve. Venting the GHe causesthe valve to then close. With the elimination of GHe from the vehicle,the IVF system provides either gaseous H2 or O2 as a replacement. Unlikein a GHe supplied system where the amount of gas is strictly limited, anIVF based system enables the actuation gas to be flowed through thevalve actuator so that the temperature conditions within the valveactuation cavity remain stable over extended durations. The performanceof the valve can be stabilized and control improved.

The IVF system can supply either one way or recirculating purges ofeither GH2 or GO2 to either prevent the ingress of external atmosphereinto components on the vehicle, maintain stable temperature conditionsat sensors or to thermally condition components such as avionics boxes,actuators, or to provide vapor cooling of structures. Vapor cooling is atechnique whereby heat is blocked from moving down a structure byintercepting it with cold gas. Such techniques can drastically reduceheating in the main vehicle tanks and further extend flight operationsby suppressing boil off. For example, a vapor retention device may belocated at the junction of a propellant tank and the feed line where theliquid propellant exits the tank. The device has apertures that permitliquid propellant to exit the tank and restrict gaseous propellant fromentering the tank through the feed line. In one instance, uponcompletion of an engine burn, the feed line between the tank and theengine is filled with liquid propellant. As the liquid furthest from thetank and closest to the engine warms, it transforms into a gas. Theretention device at the port captures rising gas bubbles at the port dueto surface tension effects. This prevents the warmer gas from enteringthe tank and warming the liquid propellant. In addition, by controllingthe movement of the gas, as liquid propellant within the feed linecontinues to transform into gas, the gas pushes the remaining liquidpropellant toward the tank, thus recapturing some of the liquidpropellant and making it available for later use. The gas forms aninsulative barrier to prevent boil off of more liquid propellant in thetank and prevent slip flow and liquid propellant from returning into thefeed line. After every burn of the engine, the process repeats.

While the IVF system takes advantage of the internal combustion engineto provide electrical power for vehicle systems and for the operation ofthe IVF pumps, this is not the boundary of what can be done with thepower produced. The nature of the IVF system is that it taps less thanthe total power which is available from the engine. The shaft power canbe used for any function in the nature of an auxiliary power unit. Itcan circulate fluids for cooling or hydraulic power and can drive muchlarger pumps than those described for the internal IVF use. Large boostpumps which raise the pressure of propellants entering the main enginescan be directly driven by the IVF engine and such pumps can also be usedto circulate propellants within the vehicle or between docked vehicles.Boost pumps can further minimize or eliminate the need for direct tankpressurization since they provide the main engine's required NPSP bydirect application of work to the fluid. The electrical generationsystem can be augmented with multiple generators including those forhigher voltage which supports the use of multiple or higher powerelectromechanical actuators for driving engine thrust vectoring or otheruses. The IVF engine can be used on an intermittent basis in concertwith solar power or fuel cell systems. These systems can be sized foraverage loads but the IVF engine can be activated when peak demands areexpected such as prior to and during main engine burns. In this waythese other systems can be reduced in mass and cost with an overallbenefit to system performance.

Moreover, the entire IVF system can be placed in a safe condition andvented of gases if it is not needed. This is a common requirement forrendezvous and docking with crewed space stations. Unlike a hydrazinesystem which can at best isolate the remaining propellant withpyrotechnic valves, the IVF can dump the accumulators and render itselfcompletely inert. The possibility of inadvertent operation of a thrusteror engine is thus completely eliminated.

In a preferred embodiment, a modular design is provided for theintegrated fluid system. More specifically all elements are mounted to asingle common panel and share a single set of fluids and electronicinterfaces to the main vehicle. The module in one embodiment wouldcontain two opposing pairs of pitch thrusters, one pair of yawthrusters, and a pair of axial or vehicle settling thrusters. Theaccumulators, being quite small, can be closely coupled to the thrustersand also the pressurization control valves with a minimum of interveningplumbing thus minimizing leak sources and component count. The internalcombustion engine, starter/generator and all electronic controls canshare a common radiation enclosure which enables thermal stabilizationof the components in space within a band near room temperature. Thecombination of all these masses on a rigid, shock and vibration isolatedpanel suppresses the movement of vibration energy both to and from themodule. The module can be completely inspected and validated prior toinstallation on the vehicle. In the event of a fault after installation,it can be readily removed and replaced as a unit. In the preferredembodiment, the module can be mounted on existing available space on theaft deck of the vehicle and because of its small size, no modificationsare required to the existing vehicle.

Regarding the second exemplary embodiment, in accordance with otheraspects of the present disclosure, the system, methods and sub-systemsor sub-combinations may be varied and still provide all required vehiclefunctions including attitude control, propellant settling, tankpressurization and venting, hardware actuation and purging and powergeneration without the need for any consumable fluids other than theliquid oxygen and liquid hydrogen present in the main vehicle tanks.

According to some aspects of the present disclosure, gas accumulatorsmay be eliminated. It can be more efficient to recirculate gas from theullage to the IVF modules and then add it back into the ullage at ahigher enthalpy. By more efficiently and effectively capturing andmoving heat, larger tanks, such as main propellant tanks for longduration flights, may be pressurized as needed. Small flow rates causegas to become super-hot. According to aspects of the present disclosure,gas flow rates may be increased, compared to other embodiments, withhigh speed electrically-powered compressors and with system pressuredifferentials. However, to keep the power demands within reason, thepressure rise from the compressors must be small. This requires thesystem to work at low pressures. For example, the thrusters and the ICEinduction must work at low pressure. According to aspects of the presentdisclosure, separate axial and lateral thrusters may be replaced withfewer low pressure thrusters that work for all attitude adjustments andsettling functions. According to aspects of the present disclosure, theIVF module may function between 10 and 100 psia. This means that theeffects of leaks are about as small as possible and weight is absolutelyminimized.

It is also an option not to use heat from the thrusters to vaporizeliquid oxygen or liquid hydrogen for increasing enthalpy. For example, amulti-cylinder internal combustion engine, such as a 6-cylinder ICE, mayreplace a Wankel engine. The ICE has the excellent balance and lowvibration of a Wankel engine, but provides larger heat rejection area tocapture heat required to do the larger scale pressurization. An ICE canalso deliver target power with at least one cylinder not firing. Evenwith reduced pressure rise, the system demands a power output on theorder of 70 HP (per module).

According to some aspects of the present disclosure, use of Nitronic 50alloy as the material of the intake valves within the ICE eliminates theneed for lubricating oil along the valve stem which would otherwise comeinto contact with pure oxygen and lead to formation of apressure-sensitive gel capable of explosive reaction.

According to further aspects of the present disclosure, independentcoolant systems may be incorporated to extract heat instead of, or inaddition to, using the vented gases from the propellant tanks to captureheat directly from heat sources. Thus, a coolant, such as Krytox, orfluorocarbon ether polymers of polyhexafluoropropylene oxide, with achemical formula: F—(CF(CF3)-CF2-O)n—CF2CF3, where the degree ofpolymerization, n, generally lies within the range of 10 to 60, may becirculated by a circulation pump in a loop including, for example, theICE, a heat exchanger associated with a compressor for pressurizing thevented gas, and the generator associated with the output of the ICE.Heat is extracted from the ICE to increase the temperature of thecoolant. In at least one embodiment, a heat exchange structure surroundseach ICE cylinder and allows higher heat transfer coefficients and hencegreater, more consistent and controlled heat extraction from thecylinders while minimizing the mass of coolant required to fill the ICEcooling channels. The coolant then interfaces with the vented gas in aheat exchanger to warm the vented gas. In at least one embodiment theheat exchanger is a counterflow centrifugal heat exchanger. The reducedtemperature coolant then cools the gas compressor motors and thegenerator before returning to interface with the ICE. Optionally, ifdesired, the coolant may also interface with the thrusters and/or othercomponents to capture further heat. An actively cooled exhaust valveassociated with the ICE enables modulation of the pressure within theexhaust manifold and hence the heat transfer to the coolant from the hotexhaust gases. This valve also maintains pressure within the ICE when itis shut down during extended duration flight in space.

The ICE exhaust system can provide more than 50% of all the waste heatthat is reclaimed for increasing system enthalpy. This coolant systemcan increase the amount of waste heat gathered from the system and allowcontrolled modulation of heat extraction by both directly extractingheat from the ICE and bypassing coolant around the ICE therebymaximizing heat extraction during pressurization and minimizing heatextraction when the system is solely generating electricity. Use of thecoolant system to scavenge waste heat from all major electrical loads,such as the compressor motors, generator units and coolant recirculationmotors makes this energy available to the vented gases used for variousfunctions. This substantially increases the amount of available energyfor pressurization as well as cooling these motors.

According to aspects of the present disclosure, heat from the ICE isextracted and moved through the system by two independently controlled,redundantly powered coolant recirculation pumps-one for the first ventedpropellant gas (e.g., oxygen) and one for the second vented propellantgas (e.g., hydrogen). By varying the coolant flowrate through the heatexchanger associated with each propellant loop, as well as by varyingthe flowrate of the coolant in each propellant loop, the heat deliveredto the two sides of the system can be controlled directly without needfor additional valves or other devices. Also positive flow can beguaranteed even with one side at very low flow such that coolantfreezing is avoided. By utilizing two pumps, one for each propellant,loss of ICE coolant flow is avoided in the case of loss of one of thepumps. Loss of one pump halts the use of that associated coolant loop,but preserves the other.

Discharge from the two compressors is supplied to the propellant tanks,the ICE and the thrusters. Different terms may be used to refer to thevented gas as the gas travels to different parts of the IVF module. Forexample, the gas may be referred to a pressurant gas when the gas withincreased enthalpy is returned to the ullage space of the tanks. Inaddition, the gas may be referred to as a reactant when the gas travelsto the ICE. Since the reactants feeding into the ICE are from acompressor, the ICE thus is effectively turbo-supercharged, allowing thedisplacement and mass of the ICE to be reduced. Lastly, the gas suppliedto the thrusters may be referred to as a propellant. The higher pressuresupplied to the thrusters enables much higher thrust output withoutincreasing the thruster size and mass.

According to other aspects of the present disclosure, a separate coolantsystem may be provided to cool the IVF system batteries and electronics.In one embodiment, a control unit comprises one or more arrays oflithium ion batteries combined with multiple computer slices. Thecoolant flows through and immerses portions of the controller unit. Thebatteries are completely immersed in a flowing reservoir of oil. The oilnot only cools the batteries, but isolates the batteries from externalvibrations. A coolant feed line transports the coolant to a heatexchanger under the power of a coolant pump. The avionics heat exchangerinterfaces with either or both reactants to cool the reactants flowingtowards the ICE.

According to some aspects of the present disclosure, to modulate massflow and heat flow the discharge out of the compressor is split. Onepassage (the “hot” passage) leads to the heat exchanger and the other(the “cold” passage) does not. A “hot” pressurization valve modulatesthe gas flow through the heat exchanger to achieve variable heatrejection. A “cold” pressurization valve effectively sets the compressorpressure ratio by restricting the flow above that required by the heatexchanger. By this means both hot gases (which has flowed through theheat exchanger) and cold gases (those bypassing the heat exchanger) areproduced. Heat rejection is thus relatively independent of compressorflowrate. The combined flow is delivered to the tank ullage forpressurization. Hot and cold gases may be bled from the mainpressurization supply line to serve other functions.

The reactants (e.g. oxygen or hydrogen) cannot be at extreme cryogenicconditions when introduced into the ICE in order to protect ICElubrication and avoid combustion produced icing. According to someaspects of the present disclosure, two stages of heat exchange areavailable to condition the oxygen and hydrogen gases flowing to the ICE.Both hydrogen and oxygen feeds for combustion are warmed by the primaryheat exchanger coolant by bleeding gas from the “hot” side of theirrespective compressor/heat exchanger. The hydrogen is further warmed byremoving waste heat from the IVF controller avionics-thus reducing thisenergy demand.

According to some aspects of the present disclosure, the coolantdischarged from the heat exchanger immediately adjacent to the motordriving the cryogenic compressors may be used to both cool the motorstator and to keep the bearings at a temperature which permits use oflubricants. This eliminates additional systems to achieve this function,such as specialized cryogenic bearings, and prolongs bearing life.

According to some aspects of the present disclosure, a high-flowcapacity, sonic-metering, low-pressure-drop, variable-position pintlestyle valve may be used to modulate the flow of vented gas dischargedfrom each compressor. These valves establish the compressor operatingpoint (combined pressure rise and flowrate) by controlling the totalflow rate. They also control the heat rejection by the heat exchanger.Preferably, the heat exchanger and control valves are designed for themaximum heat rejection conditions which correlate to the highest massflow. However these sizes would be too large under partial powerconditions. Excessive gas flow would lead to overcooling of the coolantsince the gas temperature can easily overpower the ICE heat generation.

According to some further aspects of the present disclosure, a variableposition liquid propellant atomizing mixer valve can be placed upstreamof the compressor and heat exchanger. In concert with temperatureinstruments placed at the compressor inlet and an active control system,this valve is capable of atomizing liquid and mixing the atomized liquidwith incoming vent gas even with zero differential in their totalpressures. The vaporizing liquid serves to cool the incoming vent gas toa specified target value at the inlet of the compressor. It allows thesystem to operate over a wide range of temperatures and to simplify thedesign and operation of the compressor by actively holding its inletconditions within an optimal band. This allows the system to add mass tothe vent gas so that temperatures do not become excessive. Thus, therecan be two pressurization modes—recirculation and recirculation withmass addition.

According to some aspects of the present disclosure, a counter helicalflow, coaxial-annular heat exchanger may be used to induce rotation ofthe coolant and the vent gases to improve the performance, i.e. heattransfer, of the heat exchangers. A compressor expels vent gas with arotation, which means that the gas has radial and tangential velocities.The heat exchanger has fins that preserve and induce this rotation,which stratifies the gas based on density, and thus, temperature. Sincethe gas flows through an inner space of the heat exchanger, the colder,denser gas is positioned against the interface with the coolant.Likewise, the heat exchanger induces rotation of the coolant, whichflows through an outer space of the heat exchanger, and the hotter, lessdense coolant is positioned against the interface with the ullage.Therefore, the steepest possible temperature gradient is establishedbetween the coolant and the gas, which maximizes heat transfer andimproves the performance of the heat exchangers.

According to some aspects of the present disclosure, a single gimbalplatform on each module supporting two independently controlledthrusters and the exhaust nozzle from the ICE may be used. This permitscontrol of the force vector from both thrusters and ICE exhaust. Thegimbal can rotate to produce combined settling and pitch/yaw/rolltorques or can be pointed radially outboard to produce a balanced thrustof both thrusters and ICE exhaust. The platform can be gimbaled toproduce reverse thrust to achieve spacecraft separation withoutsignificant impingement of gases on the spacecraft. This eliminates theneed for spacecraft separation springs. A failure in one thruster can becompensated for by adjusting thrust on the remaining thruster with nodegradation in vehicle control.

According to some aspects of the present disclosure, a laser-sinteredadditive manufactured structure that combines both thrusters on a moduleinto a common housing along with their propellant feedlines, may beused. The single structure acts as the structural gimbal platform. Thisapproach eliminates most plumbing and leakage points on the thrustersystem downstream of the propellant valves.

According to some aspects of the present disclosure, a single-design lowfeed pressure (<60 psia) variable thrust and mixture ratio thruster maybe used to achieve vehicle control and settling with much larger vehicleand payload mass. Use of a sonic-metering variable position pintle valvefor each thruster's propellant (e.g., oxygen and hydrogen) flow controlin concert with redundant pressure and temperature instruments upstreamof the valves and an active control system permits the thruster to adaptto widely varying inlet temperatures and pressures to produce stablethrust or to increase/decrease thrust with a turndown ratio of nearly10. This variable thrust design makes the combustion device elementsmall enough to gimbal and thus allows for the removal of all but fourthrusters from the system. This leads to a drastic reduction in valves,ignition systems etc. The variable mixture ratio thruster allows thesystem to adjust its propellant consumption to match tank boiloff(typically a very low mixture ratio) while keeping mixture ratio higherfor ignition events.

According to some aspects of the present disclosure, the thruster systempreferably separates the propellant flow control valves from thethruster combustion devices by flex lines that enable large gimbalangles and also prevent thermal soak-back from the hot combustor intothe propellant valve and the propellants immediately upstream of thevalve. This prevents rapid transients in propellant inlet conditions andenables high reliability ignition events with controlled mixture ratioand avoids transient over-temperature events from excessively highmixture ratios.

Preferably, the thruster system also uses a cryogenic temperaturecapable propellant feedline rotational slip joint at the interface ofthe thrusters and their flexible feedlines. This redundantly sealedswivel permits large rotational motion of the gimbaled thrusters withoutexcessive hose twist.

According to some aspects of the present disclosure, hot and cold bleedvalves are used to supply the thrusters with variable temperaturepropellants gases. During main pressurization events when thrusters arefiring for settling the demand from the thrusters is effectively a leakin the system that reduces energy flow back to the tanks. During theseperiods the thrusters use cold propellant which contains the leastenthalpy and hence reduces the energy drain on the system. Duringperiods when removal of energy from the system is desired, as when theICE is operating to recharge batteries, propellant is bled from the hotside of the compressor. This gas has gained heat from the coolant andhence acts to cool the ICE. Any desired combination of cold and hot gascan be produced thus permitting precision thermal control of the IVFsystem across the range of output power.

According to some aspects of the present disclosure, during high poweroperations, the exhaust valve, mixer valve and thrusters may be used asthe primary thermal management system. During periods when large amountsof electrical power are desired without tank pressurization the ICE mustreject all the heat that would be normally be scavenged and used forpressurization. According to some aspects of the present disclosure,embodiments of the IVF system can produce in excess 100 kW of electricalpower that is available for payload and vehicle functions other thanpressurization. The hydrogen mixer valve, which normally supplies only afraction of the mass that enters the compressor, is opened more fully toreduce the temperature entering the compressor to saturation levels. Thecompressor does not operate under power but this cold gas flows throughthe compressor and into the heat exchanger where the gas removes heatfrom the circulating coolant. By taking advantage of the heat ofvaporization of the hydrogen far more energy can be absorbed into thisgas before the gas is exhausted out the thrusters. No oxygen flow isrequired at the thrusters since combustion at the thruster is notrequired with warm hydrogen gas being vented to space through thethruster injector. The ICE exhaust valve is adjusted to the fully openposition to minimize pressure within the exhaust manifold and hencereduce heat transfer to the manifold cooling passages. This reduces thetotal heat that must be rejected by the coolant loop up to approximately50%.

According to some aspects of the present disclosure, the one or morerechargeable batteries having commercially available lithium ion cellsmay be used. The battery is completely immersed in fluorocarbon oil as ameans to thermally control battery cell temperature by circulating oilbetween the cells. This dense oil also creates a buoyancy effect on thecells such that approximately 60% of the mass of the cells is supportedby fluid. The oil also serves as a damper to suppress cell resonancesdue to external vibration. This enables the battery to operate withgreater reliability in space flight and to sustain higher charge anddischarge rates than would otherwise be practical with conductingcooling alone.

According to some aspects of the present disclosure, the individualcells may be joined by an interconnecting tab using a visible lightlaser welding technique. The laser spot pattern allows an increase theconductive path between the cells without increasing the gage of theinterconnect. This permits a simple construction that allows highcurrent flows with minimal voltage drop.

According to some aspects of the present disclosure, a control andcommunication system architecture may be used in which each IVF modulecontains a controller consisting of multiple physically connected butfunctionally separate chassis slices each of which has dedicatedfunctionality such as running the ICE or running the hydrogenpressurization system. According to aspects of the present disclosure,each controller contains at least one slice containing rechargeablelithium ion main vehicle batteries. The slices are joined together sothat their mass is additive to that of the batteries and the coolantcontained within so that their combined mass attenuates high frequenciesand they can share a common vibration isolation and thermal controlsystem.

According to some aspects of the invention, each of one or more slicescomprises a central processing unit (CPU) for control and processing ofsystem operation and communication. One CPU slice may be redundant toand provide backup for a second CPU slice. In addition, each slicecomprises two redundant field programmable gate arrays (FPGA) thatperform the logic functions assigned to that slice. One FPGA is primary(FPGA A) and the second (FPGA B) is redundant, but receives all of thesame communications and performs all of the same decision making as theprimary FPGA, but its output signals are suppressed. Further still, eachFPGA comprises three redundant modules. If, during operation, the outputof any module differs from the other two modules, the entire FPGA isrebooted and the secondary FPGA controls operations and communications.After the primary FPGA has rebooted, the secondary FPGA can continue tocontrol operations and communications or control can pass back to theprimary FPGA.

According to some aspects of the present disclosure, avionics thermalcontrol may be achieved by circulating fluorocarbon oil using a small,redundant, electrically driven pump to cool the avionics. Thisdielectric oil, which may also be used to cool the batteries, is used toflood the battery slices and the avionics slices. The battery cells andtheir interconnections are immersed in flowing coolant to stabilize andmoderate their temperatures even under high charge and dischargeconditions. Cold coolant flows from the pump through channels in highpower dissipation slices until it returns to the battery reservoirs andthen to a heat exchanger where cold hydrogen gas is used to remove theaccumulated heat. In this way the coldest coolant is supplied to thehigh power switches and near-room temperature return coolant floods thebatteries. The coolant is capable of remaining liquid even withcryogenic hydrogen on the other side of the heat exchanger. In this waythe controller can handle the production, rectification and control ofover 50 kW of electricity without device overheat. Approximately 5 kW ofwaste heat is dissipated under peak conditions within the small confinesof the controller assembly. Coolant channels may be formed in thechassis slices using direct laser sintering to avoid extensive externalplumbing and reducing the potential for leaks.

According to some aspects of the present disclosure, a Controller AreaNetwork (CAN) with flexible data rate (FD) bus may be used as theprimary data bus for a critical space borne system. Commands andfunctional data flow over redundant CAN FD (A &B) buses. This busminimizes conductors and is resistant to noise and interference. Datareceived by any given slice FPGA A over CAN A is shared with FPGA B andvice versa. A complete loss of one bus can thus be tolerated withoutaffecting function. The controller slices can be seen as individualredundant computers with tightly bound functions with minimal directinteraction with the adjacent slices. In this way, in one embodiment,approximately 20 computers, bound by a common redundant bus, workasynchronously to achieve IVF function with minimal central control.

According to some aspects of the present disclosure, gallium nitride(GaN) high electron mobility transistors may be used to provideradiation resistant switching in a high-power environment. GaN powerswitches are used to drive high power motors with electrical demands inexcess of 40 kW. These switches exhibit good tolerance toradiation-induced damage and minimal tendency towards radiation inducedburnout. This enables the construction of efficient, compact andcost-effective power switch inverter arrays and buck-boost electronics.

According to some aspects of the present disclosure, a high voltagepower system with variable voltage may be used to suppress coronadischarge during ascent. Corona is an electrical discharge phenomenonassociated with low pressures and a relatively high local electricalpotential. It can occur during the ascent of a rocket through the upperatmosphere as the external pressure falls. Generally voltages above 100Vare susceptible to corona depending on the gas present in thesurroundings. Helium, often used to purge vehicle systems to keepcondensable gases from freezing on cold hardware, has a relatively lowbreakdown voltage. In the second exemplary embodiment, the IVF systemoperates at two voltages: approximately 30V for typical electrical loadsand above approximately 300V for high power motors. Operation ofhardware at about 300 V during ascent would inevitably cause coronadischarge which could damage electronic components or disturb thefunction of the system. According to aspects of the present disclosure,the IVF power management slices adjust their high-Voltage power outputto approximately 100V after liftoff and keep this voltage below thebreakdown level. The power output of the system is limited during thisperiod but basic functions of the lubricant and coolant pumps as well aslow compressor flows can be supported. After vacuum is achieved voltagecan then be raised without concern about corona. The IVF system can alsocommence a purge of the controller prior to liftoff with gaseous oxygen.Oxygen has a very high breakdown voltage that is in excess of 400V.Operation with an O2 purge can be conducted at full operating voltage.

According to some aspects of the present disclosure, the internalcombustion engine and associated generator may have a maximum ratedoutput of approximately 300V. In at least one embodiment, the batterymay also produce 300V. With this configuration, the battery output issufficient to drive components with large voltage requirements, forexample a compressor, but the internal combustion engine and generatormust operate at full output to charge the battery. However, it should beappreciated that power switches operating at high voltages, such as300V, are more vulnerable to radiation damage and failure than switchesoperating at lower voltages, for example, approximately 100V or less. Itshould also be appreciated that high voltages can induce failure in acell, which can cause an entire, multi-cell battery to fail.

Alternatively, according to other aspects of the invention, lowervoltage batteries may be utilized, for example 30V batteries. In thisconfiguration, buck-boost converter is placed between the output of thegenerator and the batteries to adjust that output of the generator asneeded to protect and charge the batteries. In addition to thebatteries, many other components, including switches and valve motorsfor example, also operate at lower voltages. With lower voltage demands,operating the internal combustion engine at lower speeds, including atidle, can generate sufficient power to charge the batteries and alsoreduce wear and tear on the engine. Lower voltages also protect switchesagainst radiation damage. But if needed, the internal combustion engineand generator can be operated for periods of time at or above themaximum rated output to supply power as may be needed. As one example,the internal combustion engine can power the associated generator to runa 300V compressor. As another example, if one generator fails, a secondgenerator may be operated above the maximum rated output to provide somecompensation for the absence of the first generator. Similarly, theinternal combustion engine may be operated above its maximum ratedoutput for limited periods of time. It should be appreciated that thevoltage needs and output may vary depending upon the particular missionand that the maximum rated output may be increased or decreased as isappropriate.

According to some aspects of the present disclosure, placing ignitioncoils and high voltage leads within the ICE interior, while maintainingpositive pressurize within the ICE at all times, avoids corona dischargeand achieves coil cooling. The spark ignition system on the ICE mustfunction during ascent and the output of the ignition coil is in excessof 10,000V—well above the corona breakdown voltage. The ignition coilsand the high voltage lead leading to the spark plugs are placed withinthe camshaft cavity which receives a positive hydrogen pressure from thecrankcase ventilation system. This pressure is sufficient to preventbreakdown of these high voltage elements. This gas is recirculated bythe lubrication transport system. Gas and lubricating oil is impinged onthe ignition coils to maintain them within operating temperature.

According to some aspects of the present disclosure, a near zero-overlapintake/exhaust cam and independent oxygen and hydrogen intake runnersmay be used to prevent backfire of the ICE and assure proper cylindermixture ratio. Hydrogen and oxygen when present at the proper mixtureratio can ignite with very low ignition energy. When valves “overlap”the exhaust valve is still slightly open when the intake valve begins toopen. This means that, depending on the pressure in the exhaustmanifold, exhaust gases can also flow from there back into the intakemanifold. This means that if a combustible mixture is present at theintake port that it can light off before entering the combustion chambercausing a backfire. This is a common problem for hydrogen-burningterrestrial ICEs. It is of special concern for the IVF ICE at low idleconditions when operating on the ground. Under these conditions at theend of the exhaust stroke there are residual hot combustion gases withinthe cylinder that can backflow into the intake manifold upon the openingof the intake valve at the start of the intake stroke. This can initiatethe backfire—a highly undesirable behavior.

According to some aspects of the present disclosure, the ICE may usetuned but separate oxygen and hydrogen intake supply runners that remainseparate until a few millimeters from the intake valve. The geometry attheir intersection and the large density difference results in theaccumulation of a zone nearly pure oxygen immediately adjacent to theintake valve and another zone of nearly pure hydrogen immediatelyupstream. The area of local mixing is displaced from the intake valve sothat the ignition source is not present in that area. The ICE valvetiming reduces valve overlap to a very small value and hence no exhaustproducts can re-enter the cylinder and backflow to the intake area. Uponopening of the intake valve the two separated gas slugs are ingested andtheir passage by the intake valve creates a vigorous mixing that createsa combustible mixture within the cylinder. This assures that the ICE canbe readily started at low RPM.

According to some aspects of the present disclosure, two IVF modules arerequired per vehicle to provide the requisite redundancy and performancemargins. The configuration of thrusters is such that the thrusterswithin the two modules work in concert to achieve the proper roll, pitchand yaw maneuvers. Because the two modules are interconnected they canpreserve overall system function even if a single or multiple componentson one module are inoperative. For example, the GO2 compressor or pumpon one module can act to supply the oxygen to the thrusters on the othermodule and vice versa. It is within the scope of the present disclosurethat more than two IVF modules may be used. In accordance with methodsof the invention, a number of functions are provided within anintegrated fluid design. The methods provide various functions toinclude production of mechanical energy by an internal combustion enginethat has an output shaft, and the generation of electrical power throughan electrical starter/generator that communicates with the shaft of theengine. Electric current from the alternator may be stored in a battery.

Another function is sustained vehicle settling to drastically reducepropellant losses in the upper stage propellant tanks.

Another function includes a modular design for a plurality of thrustersthat utilize waste ullage gas, the thrusters being arranged for bothattitude and settling capabilities. The thrusters may utilize thehydrogen rich exhaust gas from the internal combustion engine or may betraditional combustion-type thrusters that burn the H2 and O2.

Various other features and advantages of the system and methods willbecome apparent from review of the following detailed description, takenin conjunction with the drawings.

BRIEF DESCRIPTION OF THE DRAWINGS

FIG. 1 is a fragmentary perspective view of part of an upper stage of aspace launch vehicle illustrating a first exemplary embodiment of an IVFmodule mounted to the aft deck of the upper stage;

FIG. 2 is an enlarged perspective view of the first exemplary embodimentof the IVF module;

FIG. 3 is a schematic diagram illustrating one aspect of the invention,namely, the provision of an internal combustion engine in the IVF systemto produce mechanical power;

FIG. 4 is a cross-sectional schematic diagram of an ICE of the presentinvention, in the form of a Wankel engine;

FIG. 5 is another schematic diagram for another aspect of the invention,namely, the provision of electrical power;

FIG. 6A is another schematic diagram illustrating yet another aspect ofthe invention, namely, provision of a thruster assembly for sustainedvehicle settling using exhaust gas from the ICE;

FIG. 6B is another schematic diagram for the aspect of FIG. 6A, butusing ullage gases for powering the thruster assembly;

FIG. 7 is a simplified schematic diagram illustrating one example ofport and starboard positioning of separate IVF modules for the upperstage of the vehicle;

FIG. 8 is a perspective view of one example construction for a thrusterassembly including a panel to which the thrusters may be mounted, alongwith hydrogen and oxygen manifolds for delivery fluids to the thrusters;

FIG. 9 is a schematic diagram illustrating one aspect of the invention,namely, tank pressurization and vent with accumulators;

FIG. 10A is a schematic diagram illustrating another aspect of theinvention, namely, venting the propellant tanks directly through thethrusters;

FIG. 10B is a schematic diagram illustrating the aspect of FIG. 10A, butventing through the ICE;

FIG. 11 is schematic diagram illustrating another aspect of theinvention, namely, accumulator replenishment;

FIG. 12 is a schematic diagram of one type of axial thruster, namely,exhaust gas thrusting;

FIG. 13 is a schematic diagram of another type of axial thruster,namely, one that combusts GH2 and GO2;

FIG. 14 is another schematic diagram illustrating basic functions of andexemplary embodiment of an IVF module;

FIG. 15 is a system schematic illustrating the primary elements withinone exemplary embodiment of an integrated fluid system and fluidconnections between the elements in the system;

FIG. 16 is a schematic view of a second exemplary embodiment of anintegrated vehicle fluid (IVF) system illustrating various componentsfor utilizing ullage from tanks;

FIG. 17 is a detailed schematic view of the second exemplary embodimentof an IVF system illustrating the physical lines between variouscomponents;

FIG. 18A is a detailed schematic view of the second exemplary embodimentof an IVF system illustrating the components and lines for pressurizingthe hydrogen tank;

FIG. 18B is a detailed schematic view of the second exemplary embodimentof an IVF system illustrating the components and lines for pressurizingthe oxygen tank;

FIG. 19 is a detailed schematic view of the second exemplary embodimentof an IVF system illustrating the components and lines for generatingelectric power;

FIG. 20 is a detailed schematic view of the second exemplary embodimentof an IVF system illustrating the components and lines for supplying thethruster assembly with propellants;

FIG. 21 is a detailed schematic view of the second exemplary embodimentof an IVF system illustrating the components and lines for cooling andheating various components;

FIG. 22 is a schematic view of the second exemplary embodiment of an IVFsystem illustrating the components and lines for cooling and heatingvarious components;

FIG. 23 is a perspective view of the second exemplary embodiment of anIVF module illustrating various components;

FIG. 24 is a further perspective view of the second exemplary embodimentof an IVF module illustrating various components;

FIG. 25A is a front elevation view of one exemplary embodiment of amixer valve for adding liquid to the vented gas;

FIG. 25B is a cross-sectional, side elevation view of the mixer valve ofFIG. 25A;

FIG. 26 is a cross-sectional view of one exemplary embodiment of a heatexchanger for transferring heat from the coolant in the coolant loop tothe vented gas;

FIG. 27A is a perspective view of one exemplary embodiment of an exhaustmanifold for an internal combustion engine;

FIG. 27B is a cross-sectional view of one exemplary embodiment of a heatexchanger associated with the exhaust manifold of FIG. 27A;

FIG. 28A is a perspective view of a thruster assembly illustratingvarious components for attitude adjustment;

FIG. 28B is a side elevation view of the thruster assembly of FIG. 25Aillustrating various components for attitude adjustment;

FIG. 29A is a perspective view of one embodiment of a controller unit;

FIG. 29B is a cross-sectional, front elevation view of the controllerunit of FIG. 27A; and

FIG. 30 is an exploded, perspective view of one embodiment of a battery;

FIG. 31 is a schematic of the power distribution system of the secondexemplary embodiment of the IVF system.

DETAILED DESCRIPTION

FIG. 1 illustrates the upper stage 10 of a space launch vehicle inaccordance with the first exemplary embodiment of the disclosure. Theouter covering or shell 12 is broken away to view the propellant tanks14, which comprise the liquid hydrogen tank 60 and the liquid oxygentank 62 with a common bulkhead separating the tanks. The aft of thevehicle includes a circumferential deck 16 that provides for mounting ofvarious system components 20 such as avionics, fluid and mechanicaldevices as well as the IVF module 30 of the present invention. FIG. 1also illustrates the main propulsion rockets 18 that are used to propelthe upper stage 10. In the Figure, the relatively small size of the IVFmodule 30 is shown. Preferably, there is an IVF module mounted onopposite sides of the aft deck 16. Twin IVF modules are able to generatemore than enough power to supply all of the upper stage systemrequirements, yet reduce overall vehicle weight by eliminating much ofthe wiring harness mass associated with traditional vehicles that usebattery power. The elevated DC voltages that can be provided by thebattery of an IVF module is also valuable for reducing EMA actuatormass. The particular vehicle 10 illustrated is a conceptual 41 tonpropellant capacity upper stage. However, the IVF module of the presentinvention can be used with any type of upper stage vehicle that has atleast some minimal space for mounting of exterior components.

Referring to FIG. 2 , an example is provided for a first IVF moduledesign. In this Figure, major structural components of the IVF moduleare illustrated to include a GO2 accumulator 34, a GH2 accumulator 36,and mounting straps 38 that can be used to mount the accumulators to aframe of the module. Lines 40 and 42 communicate with the accumulators34 and 36, and represent either vent, purge, or pressurization linesassociated with the accumulators. A housing 44 is provided for theinternal combustion engine (not shown), and a plurality of various othergas/liquid lines 50 are shown protruding from the frame for deliveringgas or liquid throughout the system. A thruster group or assembly 46 isillustrated as another component of the module having a plurality ofthrusters for settling and attitude control of the upper stage. Asshown, the thruster assembly 46 includes a pair of axial thrusters 98,two pairs of opposing pitch thrusters 94, and a pair of yaw thrusters96. A vehicle battery 48 is also illustrated and is secured to the IVFmodule, the battery 48 being charged by a generator connected to theoutput shaft of the ICE as discussed below.

FIG. 3 illustrates one aspect or concept of the present invention,namely, the provision of a small internal combustion engine (ICE) 80that is used to provide power for the upper stage systems. In apreferred embodiment, the size of the ICE 80 is approximately 200 cc,and runs at a preferred mixture ratio between 0.6 and 2.0. As shown inthe Figure, ICE 80 receives its GH2 fuel from the liquid hydrogen tank60 by vent line 64. The oxidizer, GO2, is provided by an oxygenaccumulator 34, through line 176, and metered through valve 76. Thehydrogen vent line 64 communicates with a hydrogen intake mixture valve72 that modulates the amount of hydrogen provided to the ICE. Dependingupon demand, the ICE can also receive hydrogen through a dedicatedhydrogen bleed line 66 that provides liquid hydrogen to the intakemixture valve 72. The metered amount of hydrogen is then combusted withthe oxygen within the ICE, thereby producing a mechanical output shownas shaft 82. The exhaust gas from the ICE 80 is captured in exhaust line84 that can be used for powering the axial thrusters as discussed below.The hydrogen vent line 64 would typically be used to dispose of wasteullage hydrogen gas. In the present invention, however, the waste ullagehydrogen is used to fuel the ICE. Optionally, the GH2 carried by line100 downstream of the valve 72 can be used to cool the engine exterior,maintain pressure in the crank case, and cool the internal rotor of theICE.

Referring to FIG. 4 a particular construction is provided for the ICE 80in the form of a Wankel engine. As illustrated, GO2 is provided throughline 176, while the GH2 is provided through line 100 downstream of theintake mixture valve 72. The hydrogen is first circulated in a gap 92between the engine block 90 and a cooling jacket 88. As the low pressureGH2 is circulated, it warms by heat transfer from the block 90, andfinally flows to the intake port 93. Valve 104 can be used to meter theGH2 flowing into the ICE. Once inside the engine, the hydrogen firstenters the fuel intake chamber 108. A solenoid injector valve 76 opensat the correct moment during the intake phase to inject the GO2. Thisinjector also prevents GO2 back flowing into the GH2 system, and alsocontrols the engine mixture ratio. As the rotor 114 rotates about theeccentric shaft 116, the hydrogen and oxygen are then compressed with anarea defined as the combustion chamber 110. Spark plugs 102 provide thesource of ignition for igniting the fuel within the combustion chamber110. The expansion of the gases in the combustion chamber provide themotive force for rotating the rotor 114, thus moving the combusted gasto the portion of the engine defined as the exhaust chamber 112. Thehigh temperature, GH2 rich and pressurized gas exits the exhaust port118 into the exhaust line 84. Although one will appreciate the simple,yet effective design for a Wankel engine incorporated in the IVF systemof the present invention, it shall be understood that a standard pistonengine can also be used as the ICE 80. The GO2 and GH2 are provided tothe piston engine in the same manner as illustrated for the Wankelengine. More specifically, the GO2 is provided through line 176, whilethe GH2 is provided through line 100 downstream of the intake mixturevalve 72. The hydrogen can be circulated in a gap between the engineblock and cooling jacket of the piston engine. As the low pressure GH2is circulated, it warms by heat transfer from the block, and finallyflows to a fuel intake port of the piston engine. Valve 104 can be usedto meter the GH2 flowing into the piston engine. Once inside the engine,the hydrogen is transferred to the cylinders. One or more injectorvalves can be used to inject the GO2 into the cylinders for mixing withthe GH2. Spark plugs 102 provide the source of ignition for igniting thefuel within the cylinders. The expansion of the gases upon ignitionprovide the motive force for rotating a crankshaft of the piston engine,and the combusted gases are evacuated from the cylinders to the portionof the engine defined as the exhaust chamber 112. The high temperature,GH2 rich and pressurized gas exits the exhaust port 118 into the exhaustline 84.

Referring to FIG. 5 , in another aspect of the present invention,electrical power is provided by an electrical alternator 86 that isdriven by the output shaft 82 powered by the ICE 80. The alternator 86in turn provides electrical current for charging a battery 48. For IVFsystem pumping requirements to charge the accumulators as discussedbelow, power can be supplied either from the battery 48, or power can beprovided by the output shaft 82 to a clutch (not shown) connected to theaccumulator pumps. The clutch can be engaged and disengaged to operatethe pumps. As the vehicle operates, the battery 48 will discharge duringpeak loading requirements, but will recharge during vehicle coasts,i.e., those times during which power demands are low. The use of arechargeable battery 48 removes previous restrictions on peak power andtotal available energy that was a problem with prior launch vehiclesystems in which power was limited to only battery power.

FIG. 6A is another schematic diagram illustrating another aspect of theinvention, namely, sustained settling modes provided by the thrusterassembly 46. The thruster assembly as mentioned includes a pair of axialthrusters 98 that provide settling thrust. For long duration, low thrustsettling, the high temperature, high pressure exhaust 84 can be useddirectly from the ICE to generate thrust 120. However, the thrust 120provided in this mode is limited by the peak mass flow through theengine and the allowable engine combustion temperature. Using theexhaust gas of the ICE is a very efficient method for sustainedsettling, since the ICE is normally operating to provide vehicle powerand will rarely cease to operate for any extended period of time.Therefore, there is a constant flow of exhaust gas 84 that can be usedfor providing thrust. In another vehicle settling mode shown in FIG. 6B,settling thrust can be provided directly through the GH2 ullage ventline 64 to the thruster assembly, with oxygen provided directly throughthe GO2 ullage vent line 68. These ullage gases are then combined andcombusted in combustion chamber of the thrusters. The ullage gasesprovide more than sufficient fuel and oxidation material for running theaxial thrusters.

Referring to FIG. 7 , a schematic diagram is provided showing that apair of IVF modules 30 is used, each having the same construction, andmounted to opposite sides of the vehicle 10 when looking at the vehicleouter diameter in schematic cross section. The IVF modules 30 aregenerally illustrated showing the thruster assemblies 46 having thepitch thrusters 94, yaw thrusters 96, and axial thrusters 98. The pairof IVF modules 30 provides redundancy without adding significant weight.

Referring to FIG. 8 , an example is provided for a specific thrusterassembly construction. Specifically, a panel 140 can be used to mountthe yaw thrusters 96 on one side of the panel, while the two pairs ofpitch thrusters 94 can be mounted on the other side of the panel 140. Ahydrogen manifold 142 comprises a plurality of lines and fittings forcarrying hydrogen to the thrusters, while an oxygen manifold 142 alsocomprises a plurality of lines and fittings for carrying oxygen to thethrusters. The axial thrusters 98 can also be mounted to the panel 140,or may be mounted to a separate panel. It is noted that the particularthruster panel assembly shown in the FIG. 8 can be modified to allow thethrusters to conveniently fit within the space available on the mountingstructure of the space vehicle. As compared to in the FIG. 2 , the FIG.8 shows a different, yet functional arrangement for the thrusters.

Referring to FIG. 9 , yet another concept is illustrated with respect totank pressurization for some embodiments of the invention. As shown,both the LH₂ 60 and LO2 tanks 62 have respective pressurization lines.Specifically, an oxygen pressurization line 78 pressurizes the oxygentank 62, while the hydrogen pressurization line 79 pressurizes thehydrogen tank 60. The accumulators 34 and 36 are maintained at anadequate pressure, and the tank pressurization controls 122 monitor andadjust pressurization. In this model, the accumulators supply all of thepressurization required for the propellant tanks to operate.

FIGS. 10A and 10AB illustrate an example of tank venting for someembodiments of the invention. Referring to FIG. 10A in one tank ventingmode, the propellant tanks can be directly vented through the axialthrusters 98. The ullage gases are combined and combusted in the axialthrusters. As shown, the GH2 vent line 64 and GO2 vent line 68 bothconnect to the axial thrusters. The high thrust forces that can begenerated with use of the ullage gases in this manner are very valuableto prevent vehicle shutdown caused by slosh of the LO2 and GH2. Thishigh thrust producing venting mode can be activated at any time torelieve pressure in the propellant tanks, as well as to provide ondemand, additional thrust for settling and attitude control. Referringto FIG. 10B in a low flow venting mode, the GH2 and the GO2 demands fromthe ICE engine 80 can be sufficient for relieving pressure in thepropellant tanks to maintain them in optimal pressure conditions. Thevent lines 64 and 68 provide the flow of GH2 and GO2, respectively tothe ICE 80. The operation of the ICE 80 in this low venting modeprovides continuous settling of the vehicle, and suppresses heatingwithin the tanks to prevent boil off of the propellants.

Now referring to FIG. 11 , in accordance with another aspect or conceptof the first exemplary embodiment of the present invention, accumulatorreplenishment is illustrated. One fundamental concept of accumulatorreplenishment is that the accumulators 34 and 36 must be pressurized.Accordingly, pumps 134 and 135 are provided to pressurize the lines 153and 152 that charge the accumulators 34 and 36, respectively. Drivemotors 132 and 133 drive the pumps 134 and 135. The drive motors 132 and133 may be powered by either the ICE 80, or may be electrically poweredby the battery 48. A LO2 bleed along with a GO2 vent from tank 62 arecontrolled respectively by a liquid inlet valve 148 and ullage gas inletvalve 150. In the FIG. 11 , these valves 148/150 are shown as a singleblock. These valves then meter the ullage gas or liquid oxygen throughthe pump 134 for ultimate delivery to the GO2 accumulator 34. The outletline 153 from the pump 134 carries the ullage gas/liquid oxygen in aheat exchange relationship through the thruster group 46, functioning toextract heat as necessary from one or more of the thrusters in theassembly 46. The line 153 then carries the gaseous oxygen to theaccumulator 34. The same arrangement is provided for hydrogen in whichliquid hydrogen or GH2 ullage are provided through the inlet controlvalves 149/151, the pump 135 delivers the liquid/gaseous oxygen throughoutlet line 152 and in a heat exchanger relationship with the thrustergroup 46. Line 152 then carries the gaseous hydrogen to the GH2accumulator 36. In summary, the motor driven pumps pressurize the ullageor liquid up to the necessary accumulator pressures. Liquid compressionenables high pressure requiring only low shaft power from the drivemotors 132 and 133. Heat is selectively added as needed through thethruster group 46 to thereby deliver primarily GH2 and GO2 through thelines 152 and 153, as most LH2 and LO2 will boil when coming in contactwith the thruster group 46.

Referring to FIGS. 12 and 13 , in another aspect of the invention,different types of axial thrusters are illustrated. Referring first toFIG. 12 , one example of exhaust gas thrusting is illustrated. Theexhaust 84 from the ICE 80 communicates with one or more inlet ports 160of a thruster 98. The GH2 rich exhaust gas at high temperature is thenrouted through internal passageways 162 of the thruster to the aft orrear end 164 of the thruster. At that point, the high temperature andpressurized gas is vented through one or more openings 168 into a firstsmaller chamber 168, through a nozzle or restriction 170, and then isallowed to expand within the cowl 172. The thrust is provided by theexpanding gas as it passes through the nozzle 170 into the cowl 172.Therefore, efficient means are provided for axial thrusting by simplyutilizing the exhaust gas from the ICE 80. Although the axial thruster98 is illustrated, it is also contemplated that the exhaust gas 84 canbe used to power any of the other thrusters.

Referring to FIG. 13 , another type of thruster is illustrated in whichullage GH2 is combined with ullage GO2 and then combusted to create gasexpansion and production of thrust. More specifically, one or moreullage gas inlets 180 are provided for receiving ullage GH2, such asthrough vent line 64. Similarly, oxygen can be provided through GO2 ventline 68. The GH2 flows through passageways 182 to cool the thruster, andthrough openings 184 to join the GO2 in the combustion chamber 188. Anignition source (not shown) ignites the GO2 and GH2, resulting in anexpansion of gas through nozzle 190 into the cowl 192. FIG. 13 alsoshows the heat exchange that can occur with the liquid or gaseouspropellants carried in the lines 152/153. As shown, a simple heatexchanger 198 is illustrated as a jacket that allows flow of thepropellants over the exterior of the thruster to absorb heat from thethruster. The propellants are then carried downstream to the respectiveaccumulators.

Referring to FIG. 14 , a system overview is provided showing the basicfunctions of the first exemplary embodiment of an IVF system. Ingeneral, the IVF system provides functions to include attitude control,sustained settling, tank pressurization, and a power supply. The ICE 80provides power for an alternator 86 to generate current to be stored bythe battery 48. The ICE 80 can also provide power to the drive motors132 and 133 for powering the oxygen and hydrogen pumps 134 and 135 inorder to pressurize the accumulators 34 and 36. The accumulators storeGO2 and GH2 at high pressures, and provide the source of high pressureto pressurize the propellant tanks. Tank pressurization controls 122monitor and maintain the LH₂ tank 60 and LO2 tank 62 at the properpressures. The exhaust gas 84 from the ICE 80 can be used to drive theaxial settling thrusters 98. Alternatively, ullage gas, supplementedwith liquid hydrogen under peak demands, provides sustained settlingthrust that greatly reduces losses in the tanks. The ICE 80 as well asthe settling thrusters 98 can be cooled from the waste ullage gases byfirst passing the gases in a heat exchange relationship prior tocombustion. The ICE 80 and the battery 48 work together to share powerdemands. Specifically, power boosts can be easily provided by changingthe fuel mixture ratio for the ICE in order to either more quicklycharge the battery 48 or to provide the necessary mechanical power forother vehicle systems.

Referring to FIG. 15 , a schematic system diagram is provided with amore detailed view of a plumbing schematic showing the system componentsof FIG. 14 and manner in which they are interconnected. Morespecifically, an IVF module 30 is shown with components, and the generalpiping connections between the components. The additional IVF module 30shown on the right side of the diagram within the dotted lines has thesame piping configuration as the fully illustrated IV module on the leftside of the figure, but for clarity, the piping configuration is notshown for the right side IVF module. The various vent, purge, and bleedlines/elements are illustrated as they communicate with the propellanttanks. As also discussed in part with reference to the prior Figures,these vent, purge and bleed lines include hydrogen vent 64, hydrogenpressurization 79, GH2 bleed 131, H2 purge 137, LH2 bleed 66, GO2 vent68, GO2 pressurization 78, and LO2 bleed 70.

For the axial thrusters 98, the schematic diagram shows the heatexchangers 198 that receive the pressurized gas/liquid through the lines152/153 that are pressurized by the pumps 134 and 135. Bypass valves 196allow the fluid/gas to be delivered directly to the accumulators withoutpassing through the heat exchangers 198. As shown, only one of the axialthrusters 98 communicates with the exhaust line 84 for receiving the GH2rich heated gas, while both of the axial thrusters are shown as beingcapable of operating as combustion type thrusters in which lines carrythe ullage GO2 and GH2 to the axial thrusters for combustion.

For the pitch and yaw thrusters, these are preferably combustion typethrusters, each receiving GH2 and GO2 from the accumulators as shown.Specifically, pitch thrusters 94 and yaw thrusters 96 receive GO2 fromline 176 that connects directly to the GO2 accumulator 34, and thrusters94 and 96 receive GH2 fuel from lines 174 that connect directly to theGH2 accumulator 36.

As also discussed previously, the combination of vent and bleed linesfrom the LH2 and LO2 tanks provide fuel and an oxidizer to the ICE 80that produces power for the vehicle. FIG. 15 also shows a supplementalmethod of providing oxidizer to the ICE 80 by inducting oxygen directlyinto the ICE 80 from the LO2 tank ullage instead of from the accumulator34 and through the injector 76. Specifically, FIG. 15 shows thesupplemental method by an extension of the LO2 bleed line 70 thatconnects directly to another intake port of the ICE 80. A throttle valve71 connected inline can be used to meter the LO2 into the ICE 80 at adesired rate. One advantage of this supplemental method is that the ICE80 can be operated without having to operate any system pumps.

The attitude and settling thrusters operate with combustion of thepropellants, or at least one of the thrusters can produce thrust byusing the exhaust gas from the ICE. The accumulators are pressurized,and control pressures in the propellant tanks. The IVF module is small,but can produce power and thrust to service all of the vehicles needs inthese requirements.

FIG. 16 illustrates an integrated vehicle fluid (IVF) system 210 of anupper stage of a launch vehicle in accordance with a second exemplaryembodiment of the disclosure. The IVF system 210 uses conventional gasesfrom the tanks to pressurize the tanks, generate electricity, andprovide attitude adjustment for the upper stage of the launch vehicle.According to aspects of the present disclosure, the IVF system 210comprises tanks 212 and at least one IVF module 214. In the depictedembodiment, the tanks 212 are a hydrogen tank 216 and an oxygen tank218. It should be appreciated that other propellants may be utilizedinstead of, or in combination with, hydrogen and/or oxygen including,but not limited to, methane, kerosene, nitrogen tetroxide, hydrazine,hydrogen peroxide, nitrous oxide, etc. The hydrogen tank 216 has liquidhydrogen 220 and hydrogen gas 222, and the oxygen tank 218 has liquidoxygen 224 and oxygen gas 226. The hydrogen tank 216 and the oxygen tank218 have a hydrogen gas line 228 and an oxygen gas line 230,respectively, that extend from the tanks 216, 218 to an IVF module 214.Generally, hydrogen gas 222 and the oxygen gas 226 travel through theselines 228, 230 and interact with several components in the IVF module214 to provide the pressurization, electric power generation, andattitude adjustment functions for the IVF system 210. To enhanceclarity, the gases 222, 226 vented from ullages space of the tanks maybe referred to as a pressurant gas when returning to the tanks forpressurization, a reactant when entering an internal combustion enginefor combustion, or a propellant when used in a thruster. In variousembodiments of the disclosure, the IVF module 214 operates with gaspressures between 15 and 100 psia to minimize the effect of leaks and toreduce weight. Even with low pressures, the IVF module 214 can produceapproximately 70 hp or 52.2 kW of power.

A compressor 232 draws in the vented hydrogen gas 222 and increases thepressure of the hydrogen gas 222. As a result, the hydrogen gas 222circulates through the IVF system 210, and the enthalpy, or totalenergy, of the hydrogen gas 222 is increased for the various functionsof the IVF system 210. The compressor 232 expels the hydrogen gas 222into a heat exchanger 234 that transfers heat to the hydrogen gas 222,which further increases the enthalpy of the hydrogen gas 222. Afterexiting the heat exchanger 234, the hydrogen gas 222 can travel back tothe ullage space in the hydrogen tank 216 as a pressurant gas forpressurization, travel to an internal combustion engine (ICE) 236 as areactant for combustion and electric power generation, and/or travel toa thruster assembly 238 as a propellant for attitude adjustment. Theoxygen gas 226 travels through the IVF module 214 in a same or similarmanner using a separate compressor 240 and a separate heat exchanger242.

If the hydrogen gas 222 travels back to the hydrogen tank 216, theincreased enthalpy of the hydrogen gas 222 increases the pressure withinthe hydrogen tank 216. After the launch vehicle has reached orbit andthe main engines have shut off, it may be necessary, for instance, tore-start the main engines to change the orbit or flight path of thevehicle. The pumps (not shown) that supply the main engines withpropellants have a requisite intake pressure for operation. Therefore,the IVF system 210 can pressurize the hydrogen tank 216 to supply themain engine pumps with propellant at a predetermined pressure.Similarly, the oxygen gas 226 can travel back to the oxygen tank 218 asa pressurant to increase the pressure within the oxygen tank 218.

If the hydrogen gas 222 and the oxygen gas 226 with increased enthalpytravel to an ICE 236 as reactants, the hydrogen gas 222 is combined withthe oxygen gas 226 and then combusted in the ICE 236. As shown in FIG.16 , the ICE 236 generates heat, which is scavenged by a coolant system(described below) and supplied to the heat exchangers 234, 242 toincrease the enthalpy of the hydrogen gas 222 and the oxygen gas 226.The ICE 236 produces mechanical work by rotating a shaft. A generator244 is operatively interconnected to the shaft of the ICE 236 andgenerates electric power from the mechanical work. As shown in FIG. 16 ,the generator 244 supplies electric power to a controller unit 246,which then directs the supply of electric power to the compressors 232,240, other various electric loads 248 in the IVF system 210, for examplebattery recharging, and electric loads 250 in the upper stage of thelaunch vehicle. For example, the generated electric power can be used torun a boost pump for the liquid oxygen as the main engines run to helpprevent cavitation within the liquid oxygen pump supplying the mainengines and to assist the oxygen pump.

If the hydrogen gas 222 and the oxygen gas 226 travel to the thrusterassembly 238 as propellants, the propellants 222, 226 are vented orcombusted in at least one thruster. As described in greater detailbelow, in at least some embodiments of the invention the at least onethruster is positioned on a platform that can gimbal relative to theupper stage of the launch vehicle. The at least one thruster providesattitude adjustment of the upper stage of the launch vehicle among otherfunctions such as settling liquids in the tanks 212.

While FIG. 16 illustrates the general relationship between components ofthe IVF system 210, FIG. 17 illustrates some of the physical linkagesbetween components of the IVF system 210. The IVF system 210 comprisesthe tanks 212, a first IVF module 214, and a second IVF module 252 inthis embodiment. The first IVF module 214 comprises compressors 232,240, heat exchangers 234, 242, an ICE 236, a thruster assembly 238, agenerator 244, and a controller unit 246. Though not depicted, thesecond IVF module 252 may also comprise these components in the same, orsimilar, configuration.

Fill and vent lines are shown for the tanks 212. A hydrogen fill line254 supplies liquid hydrogen 222 to the liquid hydrogen line 270 and thehydrogen tank 216, and a hydrogen vent line 258 can vent hydrogen gas220 from the ullage space of the hydrogen tank 216 and the hydrogen gasline 272. A hydrogen fill valve 256 controls the hydrogen fill line 254,and a hydrogen vent valve 260 controls the hydrogen vent line 258.Similarly, an oxygen fill line 262 supplies liquid oxygen 224 to theliquid oxygen line 290 and the oxygen tank 218, and an oxygen vent line266 can vent oxygen gas 226 from the ullage space of the oxygen tank 218and the oxygen gas line 292. An oxygen fill valve 264 controls theoxygen fill line 262, and an oxygen vent valve 268 controls the oxygenvent line 266. While the launch vehicle is on a ground surface prior tolaunch, the ambient environment constantly transfers heat to the tanks212, which boils off liquid oxygen or liquid hydrogen. The fill and ventvalves can manage the boil off by venting the gas and supplying thetanks 212 new liquid oxygen or liquid hydrogen as needed.

In addition, according to some aspects of the present invention, the gaslines 272, 292 and related pressurant lines 284, 304 can be stand pipeswithin the respective tanks 216, 218. In some embodiments, the standpipes can be between approximately 3 to 4″ in diameter and made from apolymeric composite, which is lightweight and can be buoyed within theliquid oxygen or liquid hydrogen. In addition, the stand pipes may havelow heat transfer coefficients to help insulate the liquid oxygen orliquid hydrogen from, for example, warmer vent or pressurant gas. Thiscreates temperature gradients within the tanks 212. Fiber optic orsilicon diode temperature sensing devices can utilize the temperaturegradients when the tanks 212 are in a low gravity environment to betterestimate the amount of propellant within the tanks 212.

FIG. 18A illustrates the IVF system 210 and the components thatpressurize the hydrogen tank 216. As shown, the hydrogen tank 216 hasboth liquid hydrogen 220 and hydrogen gas 222, and a liquid hydrogenline 270 draws liquid hydrogen 220 from the hydrogen tank 216 and ahydrogen gas line 272 draws vented hydrogen gas 222 from the hydrogentank 216. The liquid hydrogen line 270 extends to a hydrogen mixer valve274, which can introduce liquid hydrogen 220 to the vented hydrogen gas222 flowing through the hydrogen gas line 272. Since the temperature andpressure of the hydrogen gas 222 are not constant in the hydrogen tank216, the hydrogen mixer valve 274 can introduce varying amounts ofliquid hydrogen 220 to control and normalize the characteristics,including temperature, of the vented hydrogen gas 222 flowing into thefirst IVF module 214. Adding liquid cools the gas, and as a corollary,adding more liquid cools the gas more than when less or no liquid isadded to the gas.

Adding liquid hydrogen 220 to the vented hydrogen gas 222 can alsocontribute to the pressurization of the hydrogen tank 216. The liquidhydrogen 220 can receive heat from one or more heat sources, such as thehydrogen gas 222 or a heat exchanger, and then vaporize from a liquid toa gas. Energy is therefore transferred from the warm vented gas to thecold liquid and their streams are combined to form a gas stream ofintermediate temperature. This temperature can be controlled bymodulating the amount of liquid added. The temperature of the gasentering the compressor and the heat exchanger are thus controlled tooptimize their operation regardless of ullage gas temperatureconditions.

The vented hydrogen gas 222 enters the first IVF module 214 and is drawninto a compressor 232, which increases the pressure of the hydrogen gas222, which increases the enthalpy of the hydrogen gas 222. Afterdischarged from the compressor 232, the hydrogen gas 222 can travelthrough a hydrogen hot line 276 and through a heat exchanger 234 orbypass the heat exchanger 234 via a hydrogen cold line 278. The hydrogengases 222 in each line 276, 278 have differing enthalpies and can becombined in varying amounts to supply the hydrogen tank 216 with thedesired amount of enthalpy increase, and thus, pressure increase.

In the heat exchanger 234, coolant from a coolant loop transfers heat tothe vented hydrogen gas 222, which increases the enthalpy of thehydrogen gas 222. The heat exchanger 234 discharges the hydrogen gas222, and the hydrogen hot line 276 extends to a joint where the hydrogenhot line 276 merges with the hydrogen cold line 278. As noted above, thehydrogen cold line 278 contains hydrogen gas 222 that bypassed the heatexchanger 234 and did not receive a further increase in enthalpy.

A hydrogen hot pressurant valve 280 on the hydrogen hot line 276 and ahydrogen cold pressurant valve 282 on the hydrogen cold line 278 canmodulate the flow rate of hydrogen gas 222 through each line 276, 278 tocontrol the characteristics of the combined pressurant hydrogen gas 222after the lines 276, 278 merge together. For instance, the hydrogen hotpressurant valve 280 can open more and/or the hydrogen cold pressurantvalve 282 can close more to increase the enthalpy of the pressuranthydrogen gas 222, and vice versa to decrease the enthalpy of thepressurant hydrogen gas 222. This modulation determines the resultingpressure increase in the hydrogen tank 216.

FIG. 18B illustrates the IVF system 210 and the components thatpressurize the oxygen tank 218. The components used to pressurize theoxygen tank 218 operate in the same or similar manner as those for thehydrogen tank 216. Thus, analogous to the description for FIG. 18A, aliquid oxygen line 290, an oxygen gas line 292, an oxygen mixer valve294, an oxygen hot line 296, an oxygen cold line 298, an oxygen hotpressurant valve 300, an oxygen cold pressurant valve 302, and an oxygenpressurant line 304 operate to pressurize the oxygen tank 218.

FIG. 19 illustrates the IVF system 210 and the components that produceelectric power. In the depicted embodiment, the ICE 236 uses hydrogengas 222 and oxygen gas 226, or reactants, to turn a crankshaft 322. Agenerator 244 is operatively interconnected to the crankshaft 322 anduses the rotating crankshaft 322 to generate electric power. Thereactants 222, 226 are supplied by the hydrogen hot line 276 and theoxygen hot line 296, respectively. A hydrogen ICE line 306 extends fromthe hydrogen hot line 276, and a hydrogen ICE bleed valve 308 controlsthe flow of hydrogen gas 222 into the hydrogen ICE line 306. Similarly,an oxygen ICE line 312 extends from the oxygen hot line 296, and anoxygen ICE bleed valve 314 controls the flow of oxygen gas 226 into theoxygen ICE line 312. Reactants 222, 226 provided by the hot lines 276,296 are not combined with gases 222, 226 from cold lines that bypass theheat exchangers. Generally, the ICE 236 is more efficient when suppliedwith reactants 222, 226 that have higher enthalpy. However, it will beappreciated that other embodiments may include gases 222, 226 from coldlines to control the enthalpy of the reactants 222, 226 flowing into theICE 236.

Next, the hydrogen gas 222 travels to a hydrogen throttle valve 310, andthe oxygen gas 226 travels to an oxygen throttle valve 316. Thesethrottle valves 310, 316 precisely control the rate at which thereactants 222, 226 flow into the ICE 236. The throttle valves 310, 316can be variable-position, pintle valves with redundant brushless DCmotor linear actuators to accurately and precisely establish the flowrates of the reactants into the ICE 236. After the throttle valves 310,316, the ICE lines 306, 312 split to deliver the reactants 222, 226 toan intake manifold 318 of the ICE 236. In the depicted embodiment, theICE 236 is a six cylinder engine. Therefore, each ICE line 306, 312splits into six smaller lines to deliver the reactants 222, 226 to eachcylinder 320. Within the intake manifold 318, the reactants 222, 226combine just before entering the combustion chamber, which is defined inpart by the cylinder 320. Ignition of the combined reactants 222, 226drives a piston down the length of the cylinder 320, which turns thecrankshaft 322. Rotation of the crankshaft 322 turns the generator 244that produces electric power. On the exhaust stroke, the piston travelsback up the cylinder 320, and the combusted reactants 222, 226, or gasexhaust, are expelled through an exhaust manifold 324 of the ICE 236.

The gas exhaust travels through the exhaust manifold 324 and out into anexhaust line 326. An exhaust valve 328 controls the flow rate of theexhaust gas through the exhaust line 326 and hence the amount of backpressure that extends back through the exhaust manifold 324 and into theICE 236. Further, the exhaust valve 328 can shut off when the ICE 236shuts off to keep the ICE 236 pressurized for re-starting the ICE 236and to prevent coronal discharge. The exhaust gas in the exhaust line326 travels to an exhaust nozzle 330 of the thruster assembly 238 tocontribute to the attitude adjustment function of the IVF system 210.

When the IVF system 210 requires large amounts of electric powergeneration without tank pressurization, then the IVF system 210 mustmanage the heat generated by the ICE 236. The hydrogen mixer valve canadd liquid hydrogen to the hydrogen gas for vaporization to absorb heatgenerated by the ICE 236. Vaporization of the liquid hydrogen requiresadditional energy and allows the coolant loop to transfer additionalheat from the ICE 236 to the liquid hydrogen. Operation of thecompressor is not necessarily required, and the vaporized hydrogen gascan be vented out of the thruster assembly 238 with or withoutcombustion. In addition, the exhaust valve can open more fully such thatmore heat is expelled through the exhaust nozzle 330 of the thrustergroup 238 and less heat is transferred from an exhaust manifold of theICE 236 to the coolant loop.

Also shown in FIG. 19 is a system that ventilates the interior spaces ofthe ICE 236. While the ICE 236 is running, the combustion of reactants222, 226 can create water vapor that blows by the piston and into thecrank case. Having water vapor in the crankcase and/or other interiorspaces of the ICE 236 after the ICE 236 has stopped running isundesirable. To ventilate or purge these interior spaces, a hydrogenventilation line 332 extends from the hydrogen ICE line 306 to the ICE236, and hydrogen gas 222 travels through the hydrogen ventilation line332 and into the interior spaces of the ICE 236. The hydrogen gas 222carries the water vapor to an opposing end of the ICE 236. A hydrogenventilation line 332 can return the hydrogen gas 222 and water vapor tothe hydrogen ICE line 306. Alternatively, the hydrogen ventilation line332 can dump the hydrogen gas 222 and water vapor into the exhaust line326.

Lastly, a lubrication pump 336 is depicted in FIG. 19 . The lubricationpump 336 circulates a mixture of lubricant and blowby gases (e.g. watervapor or hydrogen) through the crankcase and other interior spaces ofthe ICE 236. The system uses the momentum of the gases to circulateliquid lubricant around the ICE without reliance on external G forces.The system can thus operate equally well under zero-G conditions andhigh G conditions. While the ICE 236 is running, the water vapor thatblows by the piston contributes to the circulation of the lubricant byimproving momentum transfer to the lubricant and hence augmenting itstransport.

FIG. 20 illustrates the IVF system 210 and the components that provideattitude adjustment for the upper stage of the launch vehicle consistentwith the second exemplary embodiment. As noted above, a hydrogen hotline 276 and a hydrogen cold line 278 have hydrogen gas 222 withdiffering enthalpies. A hydrogen hot thruster bleed valve 338 ispositioned on the hydrogen hot line 276, and a hydrogen cold thrusterbleed valve 340 is positioned on the hydrogen cold line 278. These bleedvalves 338, 340 modulate the flow of the hydrogen gases as the gasesmerge and form a combined hydrogen gas that flows through a hydrogenthruster line 342. As with the tank pressurization function, the bleedvalves 338, 340 control the enthalpy of the combined hydrogen gas tooptimize the combined hydrogen gas for use as propellants in a thruster,among other functions. For instance, use of colder hydrogen gas canreduce the enthalpy drain on the IVF system 210, and use of hotterhydrogen gas can contribute the heat rejection function of the IVFsystem 210.

Similarly, an oxygen hot line 296 and an oxygen cold line 298 provideoxygen gas with differing enthalpies, and an oxygen hot thruster bleedvalve 344 on the oxygen hot line 296 and an oxygen cold thruster bleedvalve 346 on the oxygen cold line 298 modulates the oxygen gases as theoxygen gases combine and flow through an oxygen thruster line 348.

As the hydrogen thruster line 342 approaches the thruster assembly 238,the hydrogen thruster line 342 splits into two lines, one line for eachof a first thruster 354 and a second thruster 356. A hydrogen firstthruster valve 350 controls the flow of the combined hydrogen gas intothe first thruster 354, and a hydrogen second thruster valve 352controls the flow of the combined hydrogen gas into the second thruster356. Likewise, the oxygen thruster line 348 splits into two lines, anoxygen first thruster valve 358 controls the flow of the combined oxygengas into the first thruster 354, and an oxygen second thruster valve 360controls the flow of the combined oxygen gas into the second thruster356. The mixture of propellants can be combusted in the thrusters 354,356 to provide attitude adjustment for the upper stage of the launchvehicle. Modulation of the thruster propellant valves enablesindependent control of thrust and mixture ratio for each thruster.

According to some embodiments of the present invention, the thrusters354, 356 and exhaust nozzle 330 are positioned on a common platform 362,which can be gimbaled relative to the rest of the vehicle. An elevationactuator 364 moves the platform 362 in one dimension, and an azimuthactuator 366 moves the platform 362 in another dimension. As a result,the thrusters 354, 356 and exhaust nozzle 330 can be articulated in aplurality of directions to provide attitude adjustment for the vehicle.

FIG. 21 illustrates a coolant loop 368 that scavenges heat from varioussources and transfers the heat to the hydrogen and oxygen gases throughthe heat exchangers 234, 242. The coolant loop 368 has a main coolantline 370 that circulates through the generator 244 and through the ICE236. As the main coolant line 370 travels toward the heat exchangers234, 242, the main coolant line 370 diverges into a hydrogen branch line372 and an oxygen branch line 374. The hydrogen branch line 372 travelsthrough the heat exchanger 234 for the hydrogen gas and transfers heatto the hydrogen gas. After exiting the heat exchanger 234, the hydrogenbranch line 372 scavenges heat from the motor of the compressor 232 butalso transfers heat to the bearings used within the compressor 232 toallow the bearings to function while the compressor acts on vented gasat cryogenic temperatures. Stated differently, cryogenic bearings, whichare cost prohibitive, are not required as system function is achievedwith non-specialized or customized parts, including bearings and motors.A first coolant pump 376 is positioned on the hydrogen branch line 372to circulate coolant through the hydrogen branch line 372 and,generally, the coolant loop 368.

The oxygen branch line 374 travels through the heat exchanger 242 forthe oxygen gas and transfers heat to the oxygen gas. Like the hydrogenbranch line 372, the oxygen branch line 374 exits the heat exchanger 242and scavenges heat from the motor of the compressor 240 but alsotransfers heat to the bearings within the compressor 240. The oxygenbranch line 374 has a second coolant pump 378 to circulate coolantthrough the oxygen branch line 374 and through the coolant loop 368. Thehydrogen branch line 372 and the oxygen branch line 374 merge back intothe main coolant line 370, which then extends back to the generator.

The coolant pumps 376, 378, as directed by the controller unit, cancontrol the relative amount of heat that is transferred to the hydrogengas versus the oxygen gas without the need for additional valves orother flow control devices. For instance, the first coolant pump 376 canhave a high flow rate through the hydrogen branch line 372 to transfermore heat to the hydrogen gas relative to the oxygen gas. Further, asingle pump can function to circulate coolant to prevent the coolantfrom freezing in the event that one pump fails. The coolant pumps 376,378 together also control the coolant flow through the main coolant line370, and thus, the flow rate of the coolant through the various heatsources and how much overall heat is transferred to the hydrogen andoxygen gases. Similarly, the flow rate of the gases through thecompressors and heat exchangers can be adjusted relative to the coolantflow rate to control the rate at which various IVF module components arecooled and the rate at which the gases are warmed by the oxygen andhydrogen heat exchangers 234, 242 and compressors 232, 240.

Also illustrated in FIG. 21 is a controller coolant loop 380 for thecontroller unit 246. The controller coolant loop 380 scavenges heat fromvarious components of the controller unit 246 and transfers thescavenged heat to the hydrogen gas in the hydrogen ICE line 302, whichfurther increases the enthalpy of the hydrogen gas flowing into the ICE236 to improve the efficiency of the ICE 236. The controller coolantloop 380 has a main controller coolant line 382 that flows into acontroller heat exchanger 384 to provide the additional heat transferredto the hydrogen gas.

A further illustration of the coolant loop 368 and controller coolantloop 380 is provided in FIG. 22 . In this embodiment, the sources ofheat include, the ICE 236, the generator 244, the compressors 232, 240(specifically, the motors of the compressors 232, 240), and thecontroller unit 246. One coolant loop 368 scavenges heat from the ICE236, the generator 244, and the compressors 232, 240. In the depictedexemplary embodiment, the coolant exiting the ICE 236 has a temperatureof approximately of 220° F., and the coolant is reduced to a temperatureof approximately 150° F. upon exiting the interface with the oxygen andhydrogen compressors 232, 240 and prior to interfacing with thegenerator 244. Another coolant loop 380 scavenges heat from thecontroller unit 246.

It will be appreciated that various embodiments can scavenge heat fromone or more components in various orders using any number of coolantloops. For instance, a coolant loop may circulate through the heatexchangers 234, 242 and the compressors 232, 240 before circulatingthrough an exhaust manifold of the ICE 236, through the ICE 236, andthen the generator 244 before returning to the heat exchangers 234, 242.Further, a single coolant loop can scavenge heat from all of the majorheat-producing components of the IVF module. Further still, eachcomponent may have a single coolant loop for heat scavenging. After thecoolant loop 368 passes through the exhaust manifold, the coolant loopdiverges into six separate lines, one for each of the six cylinders ofthe ICE 236. As described elsewhere herein, the lines can pass aroundthe cylinder with a particular geometry, such as a helix, to scavengeheat from the cylinder and the combustion chamber associated with thecylinder. After passing the cylinders, the six lines are collected intoa runner line, and the coolant loop 368 continues on as a single line.As an alternative, rather than having a separate coolant line for eachcylinder, multiple cylinders may be cooled by a single coolant line. Forexample, pairs of cylinders may be cooled by a single coolant line.

Also depicted in FIG. 22 is a controller coolant loop 380 that functionsto cool the controller unit 246. A coolant supply manifold or reservoir386 envelopes the controller unit 246 and batteries 388. A maincontroller coolant line 382 transports coolant from the reservoir 386 tothe controller heat exchanger 384 where the coolant interfaces with ahydrogen ICE line 306 destined for the ICE 236. The controller heatexchanger 384 warms the hydrogen gas destined for the ICE 236, therebyreducing the temperature of the coolant for purposes of cooling thecontroller unit 246. The main controller coolant line 382 returns thecoolant to the reservoir 386. A controller coolant pump 390 is part ofthe system and circulates the coolant within the main controller coolantline 382 and reservoir 386 as needed. Alternatively, the controllercoolant loop 380 could interface with oxygen gas or both hydrogen andoxygen gas.

According to aspects of the present disclosure, the spacecraft avionics,in the form of a controller unit 246, may be physically combined withthe batteries 388 for thermal control. Coolant is circulated among thebatteries 388 and controller unit 246 by small redundant, electricallydriven pumps 390. The coolant surrounding the batteries 388 andcontroller unit 246 may additionally function as a reservoir for coolantsupply. The mass of the battery cells also serves to stabilize thetemperature of the coolant and reduce variations in its temperature evenwith very small coolant mass. According to at least some embodiments,cold coolant flows through channels between individual battery cells andbetween controller slices 392 that comprise the controller unit 246. Inone embodiment, the controller slices 392 are linearly aligned with aseparate battery 388 at one end. Optionally, a second battery may bepositioned at the opposite end, or at a discrete position separate fromthe controller unit. The coolant circulation channels may pass amongindividual batteries 388 and controller slices 392 or among groups ofbatteries 388 and controller slices 392. The chassis supporting thebatteries 388 and controller slices 392 is preferably formed using lasersintering additive manufacturing. In this way, the channels and coolantflow paths are integrally formed and extensive external plumbing iseliminated thus reducing the potential for leaks.

As previously described, the controller coolant loop 380 also includes aheat exchanger 384. Heat generated by the batteries 388 and controllerunit 246 may be advantageously used to increase the enthalpy of thepropellant and power generation systems. More specifically, the heatexchanger 384 removes the accumulated heat from the batteries 388 andcontroller unit 246 which is used to increase system enthalpy and todecrease the temperature of the coolant. In this way, the coldestcoolant may be supplied to the batteries 388 and high power controllerslices 392. The coolant remains in a liquid phase even though itinterfaces with cryogenic hydrogen gas on the other side of the heatexchanger 384. Because of this cooling, the controller unit 246 canhandle production, rectification and control of over 50 kW ofelectricity without overheating. In one embodiment, approximately 5 kWof waste heat generated by the controller unit 246 and batteries 388 isdissipated in the generation of over 50 kW of electricity whilemaintaining the temperature of the coolant interfacing with thecontroller unit 246 and batteries 388 in a range of approximately 50° to70° F.

FIGS. 23 and 24 are perspective views of the first IVF module 214 thatshow the spatial relationship between various components. The hydrogengas line 272 provides hydrogen gas to the compressor 232 and the heatexchanger 234 for the hydrogen gas. Also, the oxygen gas line 292provides oxygen gas to the compressor 240 and the heat exchanger 242 forthe oxygen gas. The ICE 236 outputs mechanical work to the generator244, which transfers the mechanical work to electric power. The hydrogenpressurant line 284 returns hydrogen gas with increased enthalpy to thehydrogen tank, and the oxygen pressurant line 304 returns oxygen gaswith increased enthalpy to the oxygen tank. The controller 246 managesthe various components of the IVF system. In FIG. 24 , parts of thethruster assembly 238 are shown, including the first thruster 354 andthe gimballed platform 362.

FIGS. 25A and 25B show an exemplary embodiment of the mixer valve 274.FIG. 25A is a front elevation view of the mixer valve 274, and FIG. 25Bis a cross-sectional, side elevation view of the mixer valve 274 of FIG.25A. As stated above, the mixer valves 274 can add liquid to the gasflowing out of the tanks to control and normalize the characteristics ofthe vented gas before the gas is drawn into the compressor. The gaslines draw vented gas from the ullage space of the tanks, but in a highgravity environment, the hot returning pressurant gas may not mix wellwith the gas already in the ullage space. Therefore, the gas lines maydraw vented gas with varying characteristics. Similarly, a low gravityenvironment can result in vented gas with varying characteristics.

In the depicted embodiment, the mixer valve 274 comprises at least onespray bar 394 that extends from a central body 396, which houses liquidand a movable gate 398. With reference to FIG. 25B, the gate 398 movesslides axially within the body 396. In some embodiments, the vented gasflows in an annular space surrounding the central body 396, and thespray bar 394 extends from the central body 396 into the annular space.Each spray bar 394 has a plurality of ports 400, and liquid flows fromthe central body 396 into the spray bar 394 then exits through the ports400. In some embodiments, the liquid exits the ports 400 in a directionthat is substantially parallel to the flow of the gas. In otherembodiments, the liquid exits the port 400 in a direction that forms anangle with the flow of the gas. The angle can be between approximately 5and 45 degrees. As the gate 398 progressively opens, liquid flows fromthe central body 396 into more spray bars 394. The ports 400 may havethe same or different sized and shaped openings. The mixer valve 274 canalso include a venturi portion that accelerates the gas vented from theullage space of the tank, thus reducing the local static pressure of thegas.

In another embodiment, the mixer valve 274 draws liquid into flowing gasby way of numerous small holes or perforations in a central pintle. Theliquid is expelled at a right angle into the high speed gas stream andis immediately dispersed into droplets. A variable position gate valvevaries the amount of liquid propellant drawn into the venturi portion.As the gate progressively opens, more perforations are exposed to theventuri portion, and more small streams of liquid propellant are drawninto the venturi portion. The small diameters of the streams allow theliquid to break into small particles, which increases the surface areaof the particles for faster vaporization. The various embodiments of thehydrogen mixer valve 274 described herein can also apply to the oxygenmixer valve 294.

The compressor 214 circulates vented gas through the IVF system to servethe various functions of the IVF system. In some embodiments, thecompressor is a high-flow, high-voltage, brushless direct currentmotor-driven compressor. Regarding the pressurization function of theIVF system, since heat is added to the gas at the discharge pressure ofthe compressor, the IVF system does not require an intervening pressureaccumulator to pressurize the tanks. This also reduces the need to relyon vaporization of the liquid oxygen or liquid hydrogen to pressurizethe tanks. In addition, the pressurized gas increases the efficiency ofthe ICE for electric power generation and increases the output of thethrusters in the thruster assembly.

Valves located downstream of the compressor modulate the flow of gas andestablish the compressor operating point, which is the combination ofthe flow rate and the pressure increase of the incoming gas. Theflowrate through the hot pressurant valve also controls the amount ofheat transfer from the coolant in the coolant loop to the gas.

It will be appreciated that in some embodiments, the compressor is, forexample, a centrifugal pump or a pump that relies on rotating componentsto increase the pressure of the gas. As described below, the compressorcan discharge the gas at an increased pressure and with a rotation.Stated another way, the gas can exit the compressor having alongitudinal velocity along the line of exit, an outward radialvelocity, and a tangential velocity about the longitudinal axis of theexit line or compressor discharge. The heat exchanger can act in concertwith this rotational discharge to improve the performance of the heatexchanger.

FIG. 26 shows one exemplary embodiment of a compressor 232 and a heatexchanger 234 for the hydrogen side of the IVF module. The heatexchanger 234 transfers heat from the coolant in the coolant loop 368 tothe vented gas 222. In some embodiments, the heat exchanger 234 may be acounterflow, coaxial-annular design where each fluid flows in an annularspace with one fluid positioned outside of the other fluid. Thecompressor 232 has a motor 402 that powers an impeller 404 to draw inthe vented gas 222 and increase the pressure of the vented gas 222. Theheat exchanger 234 can function as the casing for the compressor 232.More specifically, the compressor 232 discharges the gas 222 from theimpeller through an inner space 406 of the heat exchanger 234. The innerspace 406 is in the form of a coiled tube or passageway. Coolant flowsthrough an outer space 408 of the heat exchanger. The outer space 408 isalso in the form of a coiled tube or passageway and is positionedadjacent to and in contact with the inner space 406. The rotatingimpeller 404 imparts a rotation motion on the discharged gas 222 whichis maintained as the gas travels through the coil shaped inner space406. This stratifies the gas 222 in the inner space 406 based on thedensity of the gas 222. This also stratifies the gas 222 based ontemperature as colder gas is denser than hotter gas. As a result, thecoldest gas is positioned at the interface 414 between the inner space406 and the outer space 408 of the heat exchanger 234. Similarly, theouter space 408 of the heat exchanger 234 induces a rotation motion inthe coolant flowing through the outer space 408. This rotationstratifies the coolant based on density, and thus, temperature. Sincethe coolant flows through the outer space 408, the hottest coolant ispositioned at the interface 414 between the inner space 406 and theouter space 408. As a result, the steepest possible temperature gradientbetween the gas and the coolant is established, which maximizes the heattransfer from the coolant to the gas and improves the performance of theheat exchanger 234.

In addition, though not illustrated, a bypass line can branch off afterthe impeller 404 without traveling through the inner space 406.Therefore, the vented gas in the bypass line has increased in pressurebut does not receive heat from the coolant in the coolant line. Thevarious embodiments of the hydrogen heat exchanger 234 described hereincan also apply to the oxygen heat exchanger 242.

As described above, the ICE combusts hydrogen gas and oxygen gas, orreactants, to turn a crankshaft, which powers a generator that produceselectric power. In some embodiments, the crankshaft is made oflightweight ferrium C61/C64 and the ICE is an inline six-cylinder enginemade from a laser-sintered, aluminum additive manufacturing method. As aresult of the manufacturing method, the ICE is a single block thatincludes the cylinder head, the intake manifold, and the exhaustmanifold. The single block ICE eliminates multiple critical seals ofcombustible gas and hot combustion products, which improves thereliability and performance of the ICE.

The ICE includes features that allow the ICE to perform in extremeenvironments. For example, a near zero overlap between the exhaust valveand the intake valve and independent hydrogen and oxygen intake runnerscan prevent backfiring and other issues that arise when hydrogen andoxygen are used as fuel for the ICE. A hydrogen/oxygen mixture canignite with very low ignition energy, and at the end of an exhauststroke, there are hot combustion gases in the combustion chamber. Thesehot combustion gases can flow into the intake manifold because theexhaust valve is normally still open as the intake valve begins to openfor the next stroke. Therefore, a near zero or complete zero overlapbetween the exhaust valve and the intake valve prevents the hotcombustion gases from flowing into the intake manifold, causingbackfires.

In addition, the intake runners for the hydrogen and the oxygen remainseparate until a few millimeters before the intake valve. This featurealong with the fact that hydrogen and oxygen have different densitiesresult in a volume of oxygen that is adjacent to the intake valve and avolume of hydrogen that is immediate upstream of the volume of oxygen.As a result, the area of local mixing between the oxygen and hydrogen isdisplaced from the intake valve and the hot combustion gases to preventbackfires. When the intake valve opens, the two separate volumes ofoxygen and hydrogen are drawn into the combustion chambers where thereactants are mixed and subsequently combusted.

In addition, the use of lubrication for an ICE in extreme environmentcan cause further issues. For instance, lubrication on the stem of theintake valve can interact with oxygen to form a press-sensitive gelcapable of an explosive reaction. Therefore, an intake valve thatcomprises Nitronic 50 alloy can forgo the use of lubrication and avoidthe gel-formation issue.

Embodiments of the ICE can include features that support the coolantloop to improve the ability of the IVF system to transfer heat betweenvarious components. The ICE has at least one cylinder through which apiston linearly travels. The at least one cylinder defines at least aportion of the combustion chamber, and therefore, receives heat from thecombustion process. The main coolant line of the cool loop can passthrough the ICE and through a loop around the at least one cylinder. Theloop around the cylinder can take the shape of a helix that turns uponitself. A separate coolant path may be associated with each cylinder orwith a subset of less than all cylinders. Alternatively, a singlecoolant path may traverse each cylinder.

FIGS. 27A and 27B depict an embodiment of the exhaust manifold 324 thatcollects exhaust gas from the ICE in the exhaust line and circulatescoolant from the coolant loop to scavenge heat from the hot exhaust gas.The exhaust manifold 324 can provide approximately half of the heattransferred to the coolant loop. In addition, the main coolant line canpass through, adjacent to, or otherwise receive heat from the exhaustvalve in some embodiments of the disclosure. FIG. 27A is a perspectiveview of the exhaust manifold 324 that has six exhaust channels 416, onefor each cylinder of the ICE. Exhaust gas is expelled from thecombustion chamber for each cylinder and into the exhaust channels 416.The exhaust gas then passes through exhaust manifold heat exchangers418, one for each channel 416, where the hot exhaust gas transfers heatto the coolant of the coolant loop. After the exhaust manifold heatexchangers 418, the exhaust gas is collected in an exhaust runner 418,which expels the exhaust gas to the exhaust valve and/or exhaust line.

In the depicted embodiment, coolant in a coolant channel 422 flowsthrough the exhaust manifold 324 in a direction that is opposite of theexhaust gas. The coolant enters the coolant channel 422 at the exhaustrunner 420 and, preferably then splits into six different coolantchannels 422, one for each cylinder of the ICE. Then, the coolant flowsthrough the exhaust manifold heat exchangers 418 and into the ICE whereeach coolant line further draws heat from an individual cylinder. FIG.27B is a cross sectional view of one of the exhaust manifold heatexchangers 418. As shown, the coolant channels 422 are arrayed aroundthe exhaust channel 416 so that the exhaust gas transfers heat to thecoolant. Parts of the coolant channels 422 may extend through theexhaust channel 416. Alternatively, the exhaust channel 416 may surroundthe coolant channels 422. It will be appreciated that the exhaust runner420 is also configured concentrically with the coolant channel orchannels 422 positioned around the exhaust channel 416. While theconcentric heat exchangers are utilized in this exhaust manifold 324, itwill be appreciated by those of skill in the art upon consideration ofthe present disclosure that any type of heat exchanger can be utilizedincluding, but not limited to, shell and tube heat exchangers, plateheat exchangers, plate and shell heat exchangers, adiabatic wheel heatexchangers, plate fin heat exchangers, pillow plate heat exchangers,fluid heat exchangers, waste heat recovery units, dynamic scrapedsurface heat exchangers, phase-change heat exchangers, direct contactheat exchangers, microchannel heat exchangers, etc. and any combinationthereof.

As illustrated in FIGS. 28A and 28B, first and second thrusters 354, 356are mounted to a gimballed platform 362 that permits angularrepositioning of the thrusters 354, 356 in at least two directions forattitude adjustment. An azimuth actuator 366 and an elevation actuator364 control the position of the platform 362. The lines 342, 348, 326that supply the hydrogen, oxygen and exhaust include a slip joint thatallows relative repositioning of the platform 362, thrusters 354, 356and exhaust nozzle 330. This permits control of the force vector fromboth thrusters 354, 356 and the exhaust nozzle 330. The platform 362 canrotate to produce combined settling and pitch/yaw/roll torques or can bepointed radially outboard of the spacecraft to produce a balanced thrustof both thrusters 354, 356 and exhaust nozzle 330. Further, the platform362 may be repositioned to produce reverse thrust to achieve spacecraftseparation without significant impingement of expelled gases on thespacecraft. As a result, spacecraft separation mechanisms such assprings or other devices may be eliminated. Further, a failure in onethruster may be compensated for by adjusting thrust on the remainingthruster with no loss or degradation of vehicle control.

In one embodiment, the thrusters 354, 356 are all of a single designthat operate at low feed pressure, for example, less than 60 psia, andachieve vehicle control and settling. In order to produce stable thrustor increase or decrease thrust, dependent upon then currentcircumstances, active control of the system is required. A sonicmetering variable position pintle valve is associated with the oxygenand hydrogen thruster lines 342, 348 to each thruster 354, 356 incombination with redundant pressure and temperature measurementinstruments upstream of the pintle valve provides the control. Thevariable mixture ratio provided allows the system to adjust itspropellant consumption to match tank boil off while keeping mixtureratio higher for ignition events. Thrust may be increased or decreasedin this manner, with a downturn ratio of nearly 10. This variable thrustdesign allows the combustion device element associated with eachthruster 354, 356 to be relatively small compared to thrusterspreviously used. In turn, the smaller size permits the gimbal platform362 a greater range of movement. This design also permits the reductionin the number of thrusters from previously utilized systems.

Hot and cold thruster valves and are associated with the thrusters tosupply the thrusters with variable temperature gases. During mainpressurization events when the thrusters are firing for settling, thedemand from the thrusters is effectively a leak in the energy systemthat reduces energy flow to the gas. During these periods, the thrustersuse cold gas which contains less and hence reduces energy drain on thesystem. During periods when it is desirable to remove energy from thesystem, such as when the ICE is operating to recharge batteries, gasfrom the hot side of the motor compressor unit is utilized. Any desiredcombination of cold and hot gases can be produced, thus permittingprecision thermal control of the IVF system across a range of outputpower.

After splitting, the thruster lines separate the thruster valves fromthe thrusters. Despite transporting fluid at cryogenic temperatures, thethruster lines are flexible which, in turn, permits large gimbal anglesand reduces thermal soak back from the thrusters into the thrustervalves and gas upstream of the thruster valves. This reduces rapidtransient in propellant inlet conditions and enables higher liabilityignition events with controlled mixture ratio and avoids transient overtemperature events from excessively high mixture ratios.

According to aspects of the present disclosure, the thrusters may beproduced using a laser sintering additive manufacturing process thatforms both the thrusters on a single module having a common housingalong with propellant thruster lines. This reduces and eliminatesplumbing and leakage points on the thruster system downstream of thethruster valves.

The propellant feedline rotational slip joints interface the thrusterswith flexible thruster lines. This provides a redundantly sealed swivelthat permits large rotational motion of the thrusters and exhaust nozzlewithout excessive twist of the various lines.

According to aspects of the present disclosure, a controller unit 246 isprovided with multiple controller slices 392. Each slice comprises tworedundant field programmable gate arrays (FPGA), for example, FPGA A andFPGA B. The FPGAs perform logic functions. Each slice 392 performs orcontrols one or more dedicated functions. The functions are most, if notall, spacecraft operations and could include, for example, all of thecomponents described herein, such as but not limited to powermanagement, generator output, compressor operations, internal combustionengine operations, thruster operation, gimbal control, exhaustoperation, coolant pumps, gas pumps, propellant pumps, valve operations,pressurization systems, etc. One FPGA is primary (FPGA A) and the otheris secondary (FPGA B). Both receive all data inputs and both perform alllogic functions, but the output of the secondary FPGA is suppressed.Each FPGA further comprises three redundant modules. Voting occurs amongthe three modules for each logic decision. In operation, if the outputof any one module within the primary FPGA disagrees with the output ofanother module within the same FPGA, the primary FPGA is shut down andcontrol is shifted to the secondary module and the communications of thesecondary FPGA are unsuppressed. The primary FPGA is rebooted. Oncerebooted, control is returned to the primary module and the secondaryFPGA is rebooted. In addition, one or more of the slices may comprise acentral processing unit (CPU) that operates in tandem with the otherslices in the performance of the respective functions associated withthose slices.

According to aspects of the present disclosure, an example of acontroller unit is depicted in FIGS. 29A and 29B. In the illustratedembodiment, a controller unit 246 may be physically combined with anarray of rechargeable lithium ion batteries 388, an example of which isprovided in FIG. 30 . The batteries 388 are recharged by the generator244, which is powered by the output of the internal combustion engine236. In one embodiment, the batteries 388 are directly connected to thegenerator 244, and the voltage of the batteries 388 matches the maximumrated output voltage of the generator 244. For instance, the batteries388 can have a voltage of 300V, which is also the maximum rated outputvoltage of the generator 244. Therefore, the generator 244 and therelated internal combustion engine 236 operate at a maximum rated outputvoltage and RPM, respectively, to charge the batteries 388.

In a second embodiment, the voltage of the batteries 388 is less thanthe maximum rated output voltage produced by the generator 244.Therefore, the generator 244 can charge the batteries 388 even when thegenerator 244 is producing less than maximum rated output voltage.Rather than a direct connection between the batteries 388 and thegenerator 244, a converter is positioned between the batteries 388 andthe generator 244 to step the voltage from the generator 244 down to thebatteries. In some embodiments, the maximum rated output voltage of thegenerator 244 is 300V, the output voltage of the generator 244 when theinternal combustion engine 236 is idling is 100V, and the voltage of thebatteries 388 is 30V. It will be appreciated that during emergencies,the output voltage of the generator 244, and the RPM of the internalcombustion engine 236, can increase beyond rated maximum to, forinstance, 380V.

The controller slices 392 are joined together so that their mass isadditive to that of the batteries, with the coolant flowing in channelsthrough and around the batteries 388 and slices 392. The quantity orvolume of coolant within the channels also functions as a reservoir orsupply of coolant for the cooling system. In addition to comprising athermal control system, the coolant also functions as an isolator ordampener to attenuate vibration imparted on the batteries and controllerunit by exterior sources.

According to further aspects of the present disclosure, a controllerarea network is utilized for communication and data flow among theslices 392 and the components and hardware. A schematic of powerdistribution and control is provided in FIG. 31 . Preferably, acontroller area network data bus with flexible data rates (CAN FD) isutilized. A CAN FD bus minimizes conductors and is resistant to noiseinterference. More preferably, two CAN buses (for example, CAN A and CANB) are used to provide redundancy and back up communication capability.Data received by any given slice over a first bus over the controllerarea network is shared with the slice over a second network bus. Thecommunication is received by both FPGA A and FPGA B associated with eachslice 392. A loss of one bus thus can be tolerated without affectingfunction. In effect, the controller slices 392 can be viewed asindividual redundant computers with tightly bound functions with minimaldirect interaction with the adjacent slices. In this way, for example,multiple computers (20 or more) bound by a common or redundant bus workasynchronously to achieve IVF function with minimal central control.With this architecture, delays in communication are reduced.

The power distribution in FIG. 31 also allows for redundancies inoperation in the event that a component fails. For example, thegenerator 244 is a dual stator design and a first stator 244 a supplieselectric power to a first generator slice 424 a and a first battery 388a of the controller unit 246, and a second stator 244 a supplieselectric power to a second generator slice 424 a and a second battery388 a. Each battery 388 a,b supplies a portion of the power for theremaining slices 392, a portion of the power for the hydrogen compressorslice 428, and a portion of the power for the oxygen compressor slice426. Therefore, in the event that a stator or battery fails, the otherstator and battery combination can maintain a level of functionality inthe remaining slices 392, the hydrogen compressor slice 428, and theoxygen compressor slice 426. It will be appreciated that the remainingslices 392 can include a CPU slice, a generator slice, an ICE slice, anattitude control slice, an analog IO slice, and a power managementslice.

Embodiments of the present invention may also have redundancycapabilities between IVF modules. For instance, if the batteries in oneIVF module are defective or the starter on the ICE cannot receive asufficient amount of electric power to start the ICE, then the batteryon the other IVF module can supply the requisite amount of electricpower, via a bus, to power the starter and start the ICE.

In one embodiment, gallium nitride high electron mobility transistorsare utilized in high power radiation resistant applications. Galliumnitride switches drive high power motors with electrical demands inexcess of 40 kilowatts. These switches exhibit good tolerance toradiation-induced damage and minimal tendency toward radiation-inducedburnout. This permits the construction of efficient, compact andcost-effective power switch invertor arrays and boost electronics.

Corona discharge can be a problem during ascent. Embodiments of thepresent disclosure facilitate suppression of corona discharge. Corona isan electrical discharge phenomenon associated with low pressures in arelatively high local electrical potential. Corona can occur during theascent of a rocket through the upper atmosphere as external pressurefalls. Generally, voltages above 100 volts are susceptible to coronadepending upon the gas present in the surroundings. Helium, often usedto purge vehicle systems to keep condensable gases from freezing on coldhardware, has a relatively low breakdown voltage.

In one embodiment, the IVF system operates at two voltages: 30 volts fortypical electrical loads, and above 300 volts for high power motors.Operation of hardware at 300 volts during ascent could cause coronadischarge and possibly damage electronic components or disturb thefunction of the system. The power management slices within thecontroller unit 246 adjust high voltage power output to approximately100 volts after liftoff and keep this voltage below the breakdown level.Power output of the system is limited during ascent, but basic functionsof coolant and lubricant pumps as well as low compressor flows can besupported. After vacuum is achieved, voltage may be raised withoutconcern of corona discharge. Optionally, a purge of the controller unit246 prior to liftoff may occur with gaseous oxygen. Oxygen has a highbreakdown voltage, in excess of 400 volts. With an oxygen purge,operation can be conducted at full operating voltages.

The radiation tolerance of the power switches within the controller isstrongly dependent on the voltage applied across their gate. Undernominal conditions the IVF system operates for extended durations at avery low power level. During these periods the ICE speed is held below alevel wherein the generator produces 100V instead of 300V. Thus duringthe vast majority of orbital time the primary power switches arerelatively immune to radiation-induced effects. During the peak powerdemands associated with tank pressurization or main engine poweredflight the voltage is raised to the 300V level to produce peakelectrical power when needed. During this time the probability ofradiation induced malfunction is higher but the duration at this voltagelevel is typically less than 1000 seconds. The probability of a faultduring a mission is thus minimized while still enabling multi-weekflight durations.

The spark ignition system on the ICE must function during ascent, andthe output of the ignition coil associated with the ICE is in excess of10,000 volts, well above corona breakdown voltage. To avoid or reducethe possibility of corona caused by spark ignition, the ignition coiland high voltage leads are placed within the interior of the ICE, whichspace is positively pressurized at all times. More specifically, theignition coils and high voltage leads leading to the sparkplugs areplaced within the cam shaft cavity which receives a positive hydrogenpressure from the crankcase ventilation system. This pressure issufficient to prevent breakdown of these high voltage elements. The gasis recirculated by the lubrication support transport system. Gas andlubricating oil is impinged on the ignition coils to maintain themwithin operating temperature.

The components of the IVF system can vary their operations at differentstages of a launch and flight sequence for a spacecraft. For instance,while the spacecraft is on the launch pad, the IVF system can controlthe loading and storing of propellants into the propellant tanks usingthe fill lines, the fill valves, the vent lines, and the vent valves.Due to the flow of extremely cold cryogenic fluid, typicalpneumo-mechanical vent valves exhibit crack-reseat cycling behaviorsthat result in pressure oscillations within the propellant tanks. Theuse of motor-driven pintle valves (as are used everywhere within IVF)allows the precise control of the vent flow area and hence minimizesthese pressure variations. A similar style valve can be used toprecisely control the flow of liquid propellants into and out of thetanks. With these valves and lines included within the IVF system, oncethe spacecraft is in orbit, the spacecraft can use the IVF system todock with another craft for refueling.

On the launchpad, the fill valves fill the propellant tanks with, forexample, liquid oxygen and liquid hydrogen. Once these gases displaceany inert gas within the propellant tanks, the ICE of the IVF systemstarts to bring the fluids of the IVF system, such as the coolant, up tooperating temperatures and to begin charging the batteries. This initialstart of the ICE can be on the order of hours before the launch of thespacecraft. The IVF system may also conduct a pressurization cycle totest the operation of components of the IVF system while the spacecraftis on the launchpad. The IVF system can increase the pressure in thepropellant tanks by 10 psi and then allow the pressure to drop back downto complete the test. In addition, the IVF system can test the poweroutput of the ICE, or ICEs, including a low power test of approximately20 horsepower per ICE and a high power test of approximately 72.3horsepower per ICE.

After liftoff, there are several phases that characterize IVF systemfunctions, aspects of the spacecraft flight, and the externalenvironment. The boost phase runs approximately from 0 seconds afterlaunch to 266 seconds after launch. During the boost phase, the IVFsystem minimizes high voltages as described above to mitigate coronaldischarge. In addition, the pressurization function of the IVF systemmay be reduced or stopped after the spacecraft passes a maximum dynamicpressure at a predetermined altitude since the spacecraft willaccelerate and introduce forces on the propellant tanks and the contentswithin the propellant tanks. Boost engine cut off occurs approximately266 seconds after launch, and the IVF system may need to pressurize thepropellant tanks to account for condensation within the tank. At apredetermined altitude, the external environment may cool the propellanttanks enough to cause condensation, which results in a reduction ofpressure in the propellant tanks.

Before the engines of the upper stage start, the IVF system can beginpressurizing the propellants within the tanks for use in the engines.While the engines of the upper stage propel the spacecraft, the thrusterassembly of the IVF system can assist with roll control. After theengines of the upper stage fire, or main engine cutoff, the IVF systemcan keep the ICE or ICEs running, shut off the ICEs, leave one ICErunning, etc. In some embodiments, the IVF system can run on batteriesalone for approximately 30 minutes, and the batteries can restart theICE or ICEs with approximately 15% power remaining. In sum, the variouscomponents of the IVF system are controlled to meet the pressurization,electric power generation, and attitude control functions while beingresponsive to external factors.

According to aspects of the present disclosure, the disclosed systemsare designed with redundancies to provide backup capability in the eventof component failure. For example, power from one IVF module may betransferred to another IVF module or to other spacecraft systems. if theinternal combustion engine starter is defective, batteries may be usedto start the engine.

While the present invention has been explained and illustrated withrespect to various functional features or aspects in one or morepreferred embodiments, it shall be understood that the invention can bemodified, commensurate with the scope of the claims appended hereto.Combinations and variations will occur to persons of skill in the art,all of which are deemed within the scope of the disclosure. Further, itshould be understood that each of the different concepts or aspects ofthe invention can be considered as having separate utility. Accordingly,the invention comprises a number of separate sub-combinations andcombinations that have utility with respect to supporting the functionsof an upper stage space vehicle.

What is claimed is:
 1. A system for controlling a pressure within atank, comprising: the tank storing a liquid and a gas, and the tank isassociated with an upper stage of a launch vehicle; at least one gasconduit in fluid communication with the tank; at least one liquidconduit in fluid communication with the tank; an internal combustionengine (ICE), the ICE in fluid communication with the tank to receivegas for operating the ICE; a coolant conduit having a coolant fluid, thecoolant conduit operatively associated with the ICE, wherein heatgenerated by the operation of the ICE is transferred to the coolantfluid in the coolant conduit; and a heat exchanger operativelyassociated with the coolant conduit and the at least one gas conduit,wherein heat from the coolant fluid is transferred to the gas in the atleast one gas conduit increasing the enthalpy of the gas, and at leastsome of the gas with increased enthalpy is returned to the tank toincrease the pressure within the tank.
 2. The system of claim 1, whereinat least some of the gas with increased enthalpy is transferred to theICE for operating the ICE.
 3. The system of claim 1, wherein at leastsome of the gas with increased enthalpy is transferred to at least onethruster for operation of the at least one thruster.
 4. The system ofclaim 1, further comprising: a cold gas conduit in fluid communicationwith the tank, wherein the cold gas conduit bypasses the heat exchangerto avoid receiving heat from the coolant fluid and an associatedincrease in enthalpy, and at least some of the gas without increasedenthalpy in the cold gas conduit is returned to the tank.
 5. The systemof claim 4, further comprising: a hot pressurant valve of the at leastone gas conduit that controls a flow rate of gas with increased enthalpyto the tank; and a cold pressurant valve of the cold gas conduit thatcontrols a flow rate of gas without increased enthalpy to the tank,wherein the hot and cold pressurant valves control an overall enthalpyincrease of a combined gas returned to the tank.
 6. The system of claim1, further comprising: a second heat exchanger operatively associatedwith the coolant conduit and a second gas conduit, wherein heat from thecoolant fluid is transferred to a second gas in the second gas conduitincreasing the enthalpy of the second gas, and wherein at least some ofthe second gas with increased enthalpy is returned to a second tank toincrease a pressure in the second tank.
 7. The system of claim 1,further comprising: a mixer valve operatively associated with the atleast one liquid conduit and the at least one gas conduit, wherein themixer valve is configured to introduce a liquid from the at least oneliquid conduit to the gas in the at least one gas conduit to change theenthalpy of the gas prior to the gas receiving heat from the coolantfluid in the heat exchanger.
 8. A system for controlling the enthalpy ofreactants associated with an upper stage of a launch vehicle,comprising: a first tank and a second tank associated with the upperstage of the launch vehicle; a first gas conduit in fluid communicationwith the first tank, and a second gas conduit in fluid communicationwith the second tank; an internal combustion engine (ICE) in fluidcommunication with the first and second tanks to receive a firstreactant from the first tank through the first gas conduit and toreceive a second reactant from the second tank through the second gasconduit for operating the ICE; a coolant conduit having a coolant fluid,the coolant conduit operatively associated with the ICE, wherein heatgenerated by the operation of the ICE is transferred to the coolantfluid in the coolant conduit; and a first heat exchanger operativelyassociated with the coolant conduit and the first gas conduit, whereinheat from the coolant fluid is transferred to the first reactant toincrease the enthalpy of the first reactant; and a second heat exchangeroperatively associated with the coolant conduit and the second gasconduit, wherein heat from the coolant fluid is transferred to thesecond reactant to increase the enthalpy of the second reactant, whereinat least some of the first reactant with increased enthalpy and at leastsome of the second reactant with increased enthalpy are transferred tothe ICE for operating the ICE.
 9. The system of claim 8, wherein atleast some of the first reactant with increased enthalpy is returned tothe first tank to increase a pressure within the first tank, and atleast some of the second reactant with increased enthalpy is returned tothe second tank to increase a pressure within the second tank.
 10. Thesystem of claim 8, wherein at least some of the first reactant withincreased enthalpy and at least some of the second reactant withincreased enthalpy are transferred to at least one thruster foroperation of the at least one thruster.
 11. The system of claim 10,wherein an exhaust from the ICE is expelled through the at least onethruster.
 12. The system of claim 8, further comprising: a first pintlevalve for controlling a flow rate of the first reactant into the ICE;and a second pintle valve for controlling a flow rate of the secondreactant into the ICE.
 13. The system of claim 8, wherein a portion ofthe coolant conduit extends through a portion of the ICE such that thecoolant fluid directly receives heat generated by operation of the ICE.14. The system of claim 8, wherein the first reactant is hydrogen andthe second reactant is oxygen.
 15. A system for controlling the enthalpyof a gas associated with an upper stage of a launch vehicle, comprising:a tank storing a liquid and the gas, and the tank associated with theupper stage of the launch vehicle; at least one gas conduit in fluidcommunication with the tank; at least one liquid conduit in fluidcommunication with the tank; an internal combustion engine (ICE), theICE in fluid communication with the tank to receive gas for operatingthe ICE; a coolant conduit having a coolant fluid, the coolant conduitoperatively associated with the ICE, wherein heat generated by theoperation of the ICE is transferred to the coolant fluid in the coolantconduit; and a heat exchanger operatively associated with the coolantconduit and the at least one gas conduit, wherein heat from the coolantfluid is transferred to the gas in the at least one gas conduitincreasing the enthalpy of the gas, and at least some of the gas withincreased enthalpy is transferred to at least one thruster for operationof the at least one thruster.
 16. The system of claim 15, wherein atleast some of the gas with increased enthalpy is transferred to the ICEfor operating the ICE.
 17. The system of claim 16, further comprising: asecond heat exchanger operatively associated with the coolant conduitand a second gas conduit, wherein heat from the coolant fluid istransferred to a second gas in the second gas conduit increasing theenthalpy of the second gas, and wherein at least some of the second gaswith increased enthalpy is transferred to the ICE for operation of theICE.
 18. The system of claim 16, wherein an exhaust from the ICE isexpelled through the at least one thruster.
 19. The system of claim 15,wherein at least some of the gas with increased enthalpy is returned tothe tank to increase a pressure within the tank.
 20. The system of claim15, further comprising: a mixer valve operatively associated with the atleast one liquid conduit and the at least one gas conduit, wherein themixer valve is configured to introduce a liquid from the at least oneliquid conduit to the gas in the at least one gas conduit to change theenthalpy of the gas prior to the gas receiving heat from the coolantfluid in the heat exchanger.